NASA - National Space Society · ILLUSTRATIONS Figure 1.1-l 1.1-2 1.2-l 1.2-2 1.2-3 1.2-4 1.2-5 1.2-6 1.2-7 1.2-8 1.2-9 1.2-10 1.2-11 1.2-12 1.2-13 1.2-14 1.2-15 1.3-l 2.1-l 2.1-2

NASA Contractor Report 3324

Satellite Power Systems (SPS) Concept Definition Study

Volume VII - System/Subsystem

Requirements Data Book

G. M. Hanley

CONTRACT NASs-32475 SEPTEMBER 1980

NASA

NASA Contractor Report 3324

Satellite Power Systems (SPS) Concept Definition Study

Volume VII - System/Subsystem

Requirements Data Book

G. M. Hanley

Rockwell In term tiona 2 Downey, California

Prepared for Marshall Space Flight Center under Contract NASS-32475

National Aeronautics and Space Administration

Scientific and Technical Information Branch

1980

TECH LIBRARY KAFB, NM

-. m-m.., I .m--1 m ---..n -..I. -. m-

FOREWORD

This is Volume VII - Systems/Subsystems Requirements Data Book, of the SPS Concept Definition Study Final Report as submit- ted by Rockwell International through the Satellite Systems Division. All work was completed in response to the NASA/MSFC Contract NAS8-32475, Exhibit C, dated March 28, 1978.

The SPS final report will provide the NASA with additional information on the selection of a viable SPS concept and will furnish a basis for subsequent technology advancement and verification activities. Other volumes of the final report are listed as follows:

Volume Title

I Executive Summary

II Systems Engineering

III Experiment/Verification Element Definition

IV Transportation Analyses

V Special Emphasis Studies

VI In-Depth Element Investigations

The SPS Program Manager, G. M. Hanley, may be contacted on any of the technical or management aspects of this, report. He may be reached at 213/594-3911, Seal Beach, California.

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CONTENTS

Section

1.0

2,o

3.0

4.0

SCOPE/GENEuL REQUIREMENTS . . . . . 1.1 INTRODUCTION . . . . . . . 1.2 SATELLITE POWER SYSTEM CONCEPTS . 1 .

1.2.1 Candidate Concepts . . . . 1.3 TRANSPORTATION SYSTEM . . . . . 1.4 PROGRAM GROUND RULES . . . . . FUNCTIONAL FLOW BLOCK DIAGRAMS . . . . 2.1 SATELLITE . . . . . . . .

2.1.1 Introduction . . . . . . 2.1.2 Subsystem Identification . . .

2.2 GROUND RECEIVING STATION . . . . 2.2.1 Introduction . . . . . . 2.2.2 Subsystem Identification . . .

SUBSYSTEM . . . . . . . . . 3.1 SATELLITE . . , . . . . .

3.1.1 Power Conversion . . . . . 3.1.2 Microwave Power Transmission System 3.1.3 Power Distribution and Control . 3.1.4 Structures Subsystem . . . .

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. 3.1.5 Attitude Control and Stationkeeping Subsystem. 3.1.6 Thermal Control . . . . . . . . 3.1.7 Information Management and Control . . .

3.2 GROUND RECEIVING STATION . . . . . . . 3.2.1 Rectenna . . . . . . . . . . 3.2.2 Power Distribution and Control . . . . 3.2.3 Structures . . . . . . . . . 3.2.4 Converter Stations . . . . . . . 3.2.5 Data Management and Control . . . . .

SUPPORT SYSTEMS . . . . . . . . . . . 4.1 GE0 OPERATIONAL BASE . . . . . . . . 4.2 MAINTENANCE AND REFURBISHMENT FACILITY . . . . 4.3 SPS TRANSPORTATIUN SYSTEM REQUIREMENTS . . . .

4.3.1 Transportation System Scenario . . . . 4.3.2 Heavy-Lift Launch Vehicle (HLLV) . . 4.3.3 Electric Orbital Transfer Vehicle (EOTV) .

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. 4.3.4 Personnel Transfer Vehicle (PTV)/STS-Derived

HLLV . . . . . . . . 4.3.5 Personnel Orbital Tra&fer Vehicle . . . 4.3.6 Personnel Module (PM) . . . . . 4.3.7 Intra-Orbit Transfer Vehicle (IOT;) . . .

4.4 LEO OPERATIONAL BASE . . . . . . . . 4.5 CARGO AND PERSONNEL LAUNCH AND RECOVERY FACILITIES . 4.6 BASE SUPPORT FACILITIES . . . . . . . . 4.7 LOGISTIC FACILITIES . . . . . . . 4.8 SPS GROUND RECTENNA FACILITIES . . . . . .

Page

l-1 l-1 l-2 l-2 l-18 l-19 2-1 2-1 2-l 2-l 2-10 2-10 2-10 3-l 3-l 3-l 3-8 3-17 3-22 3-25 3-31 3-33 3-43 3-43 3-46 3-50 3-53 3-53 4-l 4-l 4-l 4-l 4-2 4-6 4-10

4-12 4-15 4-18 4-18 4-19 4-19 4-19 4-19 4-19

V

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ILLUSTRATIONS

Figure

1.1-l 1.1-2 1.2-l 1.2-2 1.2-3 1.2-4 1.2-5 1.2-6 1.2-7 1.2-8 1.2-9 1.2-10 1.2-11 1.2-12 1.2-13 1.2-14 1.2-15 1.3-l

2.1-l 2.1-2 2.1-3 2.1-4 2.1-5

2.1-6 2.1-7 2.1-8 2.1-9 2.1-10 2.1-11 2.2-l

2.2-2 2.2-3 2.2-4

2.2-5

2.2-6 2.2-7 3.1-1 3.1-2 3.1-3 3.1-4

SPS Program Element Relationship . . . . Subsystem/Satellite System Relationship . . SPS Conceptual-Configuration (Nov. 1977) Solar Photovoltaic Satellite (CR-l) (Nov. i977j Solar Photovoltaic Satellite (CR-2) (Nov. 1977) Solar Photovoltaic Satellite (CR-5) (Nov. 1977) Solar Thermal Brayton (Boeing) 10 GW . . . Solar Thermal - Rankine (Nov. 1977) . . . Nuclear - Brayton (Nov. 1977) . . . . .

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. . Solar Photovoltaic Satellite (CR-2) 5 GW (Apr. 1978) . Solar Thermal - Rankine (5 GW) (Apr. 1978) . . . Microwave Transmission Subsystem - Rectenna (Apr. 1978) Rectenna Site Concept (Apr. 1978) . . . . . . NASA Reference Configuration (Oct. 1978) . . . . Alternate Satellite Concepts . . . . . . . System Efficiency Chain . . . . . . , . . Ground Receiving Station . . . . . . . . SPS Transportation System - LEO Operations Operational

. l-l

. l-2

. l-3 l-3

: l-4 . l-5

l-6 : l-7 . l-7 . l-10 . l-11 . l-11 . l-12 . l-14 . l-15

l-16 : l-17

Program . . . . . . . . . . SPS Satellite Subsystem Functi&al Relationships .

1-18 . -: 2-2

Power Generation - Photovoltaic (CR-2) . . . . . 2-2 Power Distribution . . . . . . . . . 2-3 ACS - Photovoltaic Attitude RefereAce System . . . . 2-4

Page

ACS - Photovoltaic and Solar Thermal, MW Antenna Pointing System . . . . . . . . .

- Photovol;aic Tank and Engine System . . .

ACS . . Structure - Configuration Monitor - Photovoltaic (CR-2); Thermal Requirements . . . . . . . . Microwave Antenna - Beam Generation and Control . SPS IMCS Top-Level Block Diagram . . . . . IMCS Microwave Antenna . . . . . . SPS Ground Receiving Station Sibsystem Functional

Relationships . . . . . . . . . Basic Rectenna Panel Assembly . . . . . . Panel Dipole/Diode Cluster Layout . . Ground Receiving Station Schematic Block D;agrim -

.

Preliminary . . . . . . . . . Power Conversion Station Functional Block Diagram -

Simplified . . . . . . . . . Panel Installation OperatiAns . . . . . IMCS Processor Hierarchy - Typical Groind . . . System Efficiency Block Diagram . . . . .

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. 2-12

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Simplified Integrated Block Diagram - Photovoltaic (CR-2) . 3-3 Assembly Tree - Solar Photovoltaic Power Conversion . . 3-4 Alternate Solar Cell Design . . . . . . . . 3-5

Vii

. . , @.- . . . ..--..---...--.... . . . . . -.-. ._. .._ . . . . . _ . . .._ . .__...., ,._.... ..,,... ..-.._,. .-.. ._ ._. . ,, _._._.__.__ _ -.-_

Figure

3.1-5 3.1-6 3.1-7 3.1-8 3.1-9 3.1-10 3.1-11 3.1-12 3.1-13 3.1-14 3.1-15 3.1-16 3.1-17 3.1-18 3.1-19 3.1-20 3.1-21 3.1-22 3.1-23 3.1-24

3.1-25 3.1-26 3.1-27 3.1-28 3.1-29 3.1-30 3.1-31

3.2-l 3.2-2 3.2-3 3.2-4 3.2-5 3.2-6 3.2-7 3.2-8 3.2-9 3.2-10 3.2-11 3.2-12 4.3-l 4.3-2 4.3-3 4.3-4 4.3-5 4.3-6 4.3-7 4.3-8 4.3-9 4.3-10 4.3-12

GaAlAs Solar Cell Voltage and Current Characteristics . GaAlAs Solar Cell Blanket Cross Section . . . . . Solar Panel Power Output - Watts/m2 . . . . . . Functional Block Diagram . . . . . . . . . Satellite Antenna Array Assembly . . . . . . . Assembly Tree - Microwave Power Subsystem . . . . Klystron Subarray Assembly . . . . . . . . Heat Radiators on Array Face . . . . . . . . Reference Phase Distribution System . . . . . . Klystron Power Requirements (Preliminary) , . . . Transistor MIC Amplifier . . . . . . . . . Transistor Power Circuitry . . . . . . . . Transistor Chip Layout . . . . . . . . . Power Distribution - Simplified Block Diagram . . . Assembly Tree - Power Distribution and Control Subsystem . Assembly Tree - Structures Subsystem . . . . Structure Breakdown . . . . . . . . Functional Flow Diagram . . . . . . . Satellite Coordinate Systems . . . . . . Assembly Tree - Attitude Control and Stationkeeping

Subsystem . . . . . . . . . . Thermal Control Functional Flow Diagram . . . Klystron Radiator Configuration . . . . . Assembly Tree - Thermal Subsystem . . . . IMCS - MW Antenna . . . . . . . . IMCS - Attitude Control , . . . . . . IMCS - Power Distribution . . . . . . Assembly Tree - Information Management and Control

Subsystem . . . . . . . . . Operational Ground ReLeiving Facility (Rectenna) . Ground Receiving Station Subassembly Relationships Assembly Tree - Rectenna . . . . . . . Simplified Schematic - Rectenna . . . . . Rectenna Systems Major Assembly/Component . . Rectenna Schematic Block Diagram - Preliminary . Assembly Tree - Power Distribution and Control . Rectenna Panel Assembly and Installation . . . Assembly Tree - Structures . . . . . . Rectenna Array Support Structure . . . . . Simplified Block Diagram - Converter Station . . IMCS Hierarchy - Ground Receiving Station . . SPS LEO Transportation Operations . . . . SPS GE0 Transportation Operations . . . . Reference HLLV Launch Configuration . . . HLLV First Stage (Booster) - Landing Configuratio; HLLV Second Stage(Orbiter) - Landing Configuration Selected EOTV Configuration . . . . . . LO2/LH2 SSME Integral Twin Ballistic Booster . . STS HLLV Configuration . . Liauid Rocket Booster Main Engine'(SS&-3;) . . POTV Configuration . . . . . . Advanced Space Engine . , . . . .

viii

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Page

3-5 3-6 3-6 3-9 3-10 3-11 3-11 3-12 3-12 3-13 3-14 3-14 3-15 3-19 3-20 3-24 3-24 3-27 3-28

3-31 3-32 3-33 3-34 3-37 3-38 3-39

3-40 3-43 3-44 3-44 3-45 3-46 3-48 3-48 3-51 3-52 3-52 3-53 3-54 4-2 4-3 4-7 4-8 4-9 4-10 4-13 4-14 4-15 4-16 4-17

TABLES

Table

1.2-l Solar Photovoltaic Weight Summary . . . . 1.2-2 Solar Thermal Weight Summary . . . . . 1.2-3 Nuclear Reactor Concept Weight Summary . . . 1.2-4 Photovoltaic (CR-2) Satellite Mass Statement -

1.2-5

1.2-6 1.2-7 1.4-l 1.4-2 3.1-l 3.1-2 3.1-3 3.1-4 3.1-5 3.1-6 3.1-7 3.1-8 3.1-9 3.1-10 3.1-11 3.1-12

Point Design . . . . . Solar Thermal Satellite Mass Statemeit -

:. .

Point Design . . . . . . . . . NASA Reference Satellite Mass Properties . . Mass Properties - Alternate Concepts . . . Program Ground Rules . . . . . . . General Requirements Describing Overall SPS Program . Solar Array Functional Requirements . . . . . GaAlAs Solar Cell and Blanket Preliminary Specification . SPS Reflector Preliminary Specification (CR-2) . . Solar Array Interfaces . . . . . . . . Microwave Antenna - Operating Modes . . . . . Functional Requirements . . . . . . . . Design and Performance Characteristics . . . . Phase Error Budget . . . . . . . . . Array Characteristics . . . . . . . . Klystron Power Module . . . . . . . . Transistor . . . . . . . . . . . Power Distribution and Control Subsystem -

3.1-13 3.1-14 3.1-15 3.1-16 3.1-17 3.1-18 3.1-i9 3.1-20 3.1-21 3.1-22 3.1-23 3.1-24

Operating Modes . . . . . . . . . Design and Performance Characteristics . . . . Structural Subsystem - Operating Mode . . . . Design and Performance Characteristics . . . . Attitude Control and Stationkeeping - Operating Modes Attitude Control Requirements . . . . . . Attitude Control RCS Requirements . . . . . Stationkeeping RCS Requirements . . . . . . Electric Thruster Requirements . . . . . . Thermal Control Subsystem - Operating Modes . . . Klystron Cavity Radiators . . . . . . . IMCS - Operating Modes . . . . . . . . . Preliminary Data Interface Summary - Photovoltaic

3.1-25

3.1-26 3.1-27 3.2-l 3.2-2 3.2-3 3.2-4 3.2-5

(CR-2) Configuration . . . . . . . . Preliminary Control Interface Summary - Photovoltaic

(CR-2) Configuration . . . . . . . . Hardware Summary . . . . . . . . . Weight/Power/Volume Summary - IMCS . . . . . Rectenna Functional Requirements . . . . . Rectenna Preliminary Specifications . . . . . Power Distribution and Control - Operating Modes . Design and Performance Characteristics . . . . Structural Subsystem - Operating Mode . . . .

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IX

Page

l-8 l-8 l-9

l-12

l-13 l-14 l-17 l-19 l-19 3-2 3-7 3-8 3-9 3-10 3-15 3-16 3-16 3-16 3-17 3-17

3-18 3-21 3-23 3-26 3-27 3-28 3-29 3-30 3-30 3-32 3-35 3-36

3-40

3-40 3-42 3-42 3-45 3-47 3-47 3-50 3-51

Table Page

4.3-l 4.3-2

4.3-3 4.3-4 4.3-5 4.3-6 4.3-7 4.3-8 4.3-9 4.3-10 4;3-11 4.3-12 4.3-13

TPU Transportation Requirements SPS Program Transportation Req&emeits

Construction Phase ,'30-Gear'

Total Transportation Require&n&, HLLV Sizing -

Li-Yek P;ogrk Ground Rules/Assumptions . , .

HLLV Mass Properties HLLV Weight Statement

l . . L . ,

. . . . . , HLLV Propellant Weight Sumwary . . . . . EOTV Sizing Requirements . . . l .

EOTV Thruster Characteristic; . . l .

EOTV Weight/Performance Summary , . 1 . . POTV Weight Summary Current ASS Engine Weight

. . l . . .

IOTV Weight Summary , . . . , . . . . . . L

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. l 4-4

. . 4-5 l . 4-6 . . 4-7 . . 4-8 . . 4-9 . . 4-11 . . 4-11 . . 4-12 . l 4-16 . . 4-17 . . 4-18

X

1.0 SCOPE/GENERAL REQUIREMENTS

1 .O SCOPE/GENERAL REQUIREMENTS

1.1 INTRODUCTION

This volume of Satellite Power Systems (SPS) Concept Definition Study final report summarizes the basic requirements used as a guide to systems analysis and is a basis for the selection of candidate SPS point design(s). Initially, these collected data reflected the level of definition resulting from the evaluation of a broad spectrum of SPS concepts. As the various concepts matured these requirements were updated to reflect the requirements identified for the projected satellite system/subsystem point design(s). Earlier studies (reported in Volumes I - VII, SD 79-AP-0023, dated April 1978) established two candidate concepts which were presented to the NASA for consideration. NASA, in turn, utilizing these and other concepts developed under the auspices of other contracts, established a baseline or reference concept which was to be the basis for future evaluation and point design. This volume defines the identified subsystem/systems requirements, and where appropriate, presents recommendations for alternate approaches which may represent improved design features. A more detailed discussion of the selected point design(s) will he found in Volume II of this report.

Figure 1.1-l establishes the relationship of the satellite system with the other elements of the SPS program.

Figure

. HLLV l SlS l CON

.POTV

1.1-l. SPS Program Element Relationship

l-1

Figure 1.1-2 identifies the various satellite subsystems and the functions as they apply on the satellite. Equivalent functions are applicable to the ground receiving station (rectenna) system and will not be expanded here. A limited discussion of ground receiving station subsystem functions will be found in the section dedicated to ground receiving station requirements.

SATELLITE SYSTEM

r ---- 1 GROUND 7 1 SYSTEM I L -2-- J

I ---a-)

!

l SOLAR CELLS

aSOLAR BLANKET

*REFLECTOR

*HEAT REJECTION’

l SECONDARY STRUCTURE*

*DC-RF CONV l FWR OISTRIB

*ANTENNA *REGULATION

*ANTENNA GIMBAL CONTROL’

*ENERGY STORAGE

*SECONDARY STRUCTURE

*POWER CONOITIONING

*ROTARY JOINT

*CONTROL

.SECONOARY STRUCTURE*

*CENTRAL GIRDER

4ECONOARY

.REEFLECTOR

*ANTENNA

.ROTARY JOINT

*SOLAR ARRAY

*MECHANISMS

l AlTlTUOE OETERMIN

l MW ANTEN- NA FIG. CONTROL

l CONCEN-

EToR CONTROL

l REACTION CONTROL

.POWER *PROCESSOR *RF-OC CONVERSION/ CONVERSION PHOTOVOLTAIC l ;EM;ONTROL

-POWER *HEAT OlSTRlBUTlONl

REJECTION *DATA BUS CONTROL’

*ANTENNA SYSTEMS

*REMOTE ACOUISITION &CONTROL

4NFORYATION MANAGEMENT’

%ELL .

l STRUCTURE’

*ROTARY JOINTS

*SECONDARY STRUCTURES*

*SUBMULTI- PLEXOR

.MICRO- PROCESSOR

*CONTROL*

*SECONDARY *UTILITY GRNO

STRUCTURE* INTERFACE’

. SECONOARY STRUCT*

*MAJOR INTERFACES PROCESSING’

Figure 1.1-z. Subsystem/Satellite System Relationship

The ground receiving station is identified in Figure 1.1-2 and is shown to have an indirect (dotted) relationship to the orbiting satellite. Major assemblies comprising each subsystem are identified. Unique factors such as elements of one subsystem that are integrated with another (for example, thermal radiators, subsystem control, etc.) are also identified. This document will also identify supporting subsystems, including the transportation system and SPS related ground facilities where these elements have been identified and evaluated.

1.2 SATELLITE POWER SYSTEM CONCEPTS

1.2.1 CANDIDATE CONCEPTS

Initial Candidate Concepts

Many candidate system concepts have been considered since the inception of this study program. Six satellite concepts were identified for consideration at a briefing in November 1977. These concepts are shown in Figure 1.2-1. A single rectenna farm concept was assumed, applicable to all satellite concepts,

l-2

Figure 1.2-l. SPS Conceptual Configuration (Nov. 1977)

Solar Photovoltaic (CR-l). Figure 1.2-2 illustrates the solar photovoltaic (CR-l) satellite power system concept. The CR-1 system was a planar array and had an overall planform dimension of 2.0x28.58 km. The depth of the satellite was 1.17 km. This system required 48.99 km2 of deployed solar cell area and had a total mass of 37.3~10~ kg, including a 30.7 percent growth factor. The major advantages of the CR-1 configuration were its simplicity of design; it did not require reflectors; and its relative insensitivity to misorientation angles of as much as 43 degrees. The CR-1 configuration would have had the largest solar cell area and mass in orbit.

Figure 1.2-2. Solar Photovoltaic Satellite (CR-l) (Nov. 1977)

l-3

Solar Photovoltaic (CR-2). Figure 1.2-3 illustrates the solar photovoltaic (CR-2) satellite power system concept. The CR-2 system used reflector membranes to concentrate solar energy on the cells. The satellite had two "Vee" troughs per wing. The overall planfonn dimensions were 2.75x27.16 km, and the depth was 1.2 km. This system required 23.76 km2 of deployed solar cell area and had a total mass of 33.7~10~ kg, including a 30-percent growth factor. The major advantages of the CR-2 configuration were the reduced requirement for solar cells and low weight which reduced overall cost. The disadvantages were the planform of the satellite was higher than for CR-1 and the system was sensitive to misorientation. A +l degree misorientation of the solar array required an additional 7.9 percent of reflector surface area. The reflective membranes for the GE0 environment was not available, and reflectivities of 90 percent at the beginning of life and 72 percent at the end of life were used in the design.

Figure 1.2-3. Solar Photovoltaic Satellite (CR-Z) (Nov. 1977)

Solar Photovoltaic (CR-5). Figure 1.2-4 illustrates the solar photovoltaic (CR-S) satellite power system concept. The CR-5 system had two troughs per wing and used reflector membranes to concentrate solar energy on the cells. The satellite had the overall planform dimension of 3.12x32.84 km and the depth was 1.4 km. This system required 10.4 km2 of deployed solar cell area and had a total mass of 37.4~10~ kg, including a 31.2-percent growth factor. The CR-5 system required the lowest solar cell area. The CR-5 configuration was very sensitive to misorientation angles of only rl degree. At a geometric concentration ratio of 5, an increase in reflector surface of 43 percent was required to compensate for a misorientation of kl degree.

l-4

Figure 1.2-4. Solar Photovoltaic Satellite (CR-51 (Nov. 1977)

Solar Thermal - Brayton. Figure 1.2-5 shows one Boeing concept for a lo-GW solar thermal SPS. It used four concentrator modules, each composed of thousands of planar facets which reflect sunlight into a cavity absorber. Ceramic tubing in the absorber heated pressurized helium to 1379'C (2514'F) which was supplied to a Bray:on cycle power module comprised of a turbine, regenerator, cooler, compressor, and electrical generator. Heat rejected from the cycle was dissipated by means of a NaK loop to a heat pipe/fin radiator. Microwave power was transmitted from a single antenna at the end of the satellite.

Solar Thermal - Rankine. Figure 1.2-6 shows a Rockwell concept for a 5-GW solar thermal SPS using a cesium Rankine cycle. The two concentrators were inflatable, using aluminized plastic film with a transparent canopy. Sunlight was concentrated on an open-disc absorber (cesium boiler) which provides cesium vapor at 1038'C (1900'F) to cesium turbines. Exhaust from the cesium turbines was condensed at 400'F in a tube/fin radiator. Each of the concentrator modules was hinged to permit seasonal tracking of the sun without imposing gravity gradient torques on the satellite. The beam connecting the two modules was offset to locate the rotary joint at the satellite center of gravity.

Nuclear - Brayton. Figure 1.2-7 illustrates the nuclear-powered satellite power system concept. The nuclear Brayton powered SPS consisted of 26 power modules configured in a circular array 2 km in diameter. The antenna was separated by a distance of 3 km from the power modules. In this manner, reactor shadow shielding and reactor-close plane separation distance were combined to reduce nuclear radiation at the antenna to a level acceptable to maintenance personnel. Each power module generated 344 MW, to provide the required power at the distribution bus as well as its own housekeeping requirements. Brayton cycle waste heat was rejected by a square radiator measuring 227 meters (750 feet) on each side (26 required). Each power module was approximately 40 feet in diameter and 144 feet in length, and contained one nuclear reactor with

l-5

- 8 tN KWKW,

.SECONO GENERATION NAK HEAT PIPE IAOIATORS TllRIOCOYPRESSOA l/O

PIE SC lYY/w4 ,, .;g, I f I ow IIJ&md I,, - 818 l 3M UWa POWERSAT YOOULE

IIIGH PRESSURE DESIGN

.NaK HEAT )ItE KADIATOI

Figure 1.2-5. Solar Thermal Brayton [Boeing) 10 GW

l INfl.ATAWE *FACETED -TENSIONING llEOYTS l flGUM CONlllOL I

Figure 1.2-6. Solar Thermal - Rankine (Nov. 1977)

Figure 1.2-7. Nuclear - Brayton (Nov. 1977)

shadow shield, fuel reprocessing assembly, and two closed Brayton cycle power conversion units. The power module could be removed from the radiator for replacement by remotely operated equipment.

Satellite Mass Properties. Tables 1.2-l through 1.2-3 present the summary weights for the six initial candidate satellite concepts. The solar thermal weight summary illustrates the known weight elements for both potassium-(K) and cesium-(Ce) based Rankine thermal cycles.

l-7

Table 1.2-l. Solar Phqtovoltaic Weight Summary (GaAlAs Solar Cells) (Nov. 1977)

CR.5 CR.2 CR-1 GROW” CONCENTRATlON RAT10 10s KG 106 KG l@ KG --- I

COLLECTOR ARRAY (NON ROT., I103531 PRIM.,SEC. STRUCTIMECH. 9.300 ATTITUDE CONTROL .3m SOLAR CELLS 3.297 REFLECTORS 2.182 POWER CONOIT. 307 Wl”E HARNESS/SLIP RING 2.091

“::2 2.un 0.212 ,375 5.990 12.343 2.052 N/A 307 ,387

2.891 2.469

m.31 25.0 30.0 24.7 15.0 50.0 97.0

ANTENNA l”OTATlNG, 19.7941 9.794 9.794 (23.11 PRIM /XC. STRUCT.IMECH. .268 I 250 CO”LlNG ,200 50.0 PWR COXVERTERS 5.690 20.0 WlRlNGr’SLlP RING 3% SAME SAME UO WAVEGUIDES 2U.O

IMS EOMTKABLING ,240 ,240 .24a 00.0

PROPELLANTIYEAR .100 ,103 .lcu 0

SUBTOTAL SATELLITE SYST. 28.497 25.599 29.513

GROWTH ALLOWANCE 0.002 0.115 8.754 31.2

TOTAL SATELLITE SYST. 37.379 33.714 37.270

COMPARABLE SILICON CR - i WEIGHT - 43.589 X l@ KG

Table 1.2-2. Solar Thermal Weight Summary (Nov. 1977)

CONVERSION CONCEPT

COLLECTOR ARRAY ,NON.NOT,

PRIM &EC. STRUCTJMECH.

ATTITUDE CONTROL

SOLAR COLLECTOR

SOLAR ABSORBER

TURBO EOUlP ,GEN

POWER CONDIT

WlRE HAHNESS,SLIP RING

RADIATORS

ANTENNA IROT I

IMS EGMKAOLING

PROPELLANTIYEAR

SUBTOTAL SATELLITE SY-

GROWTH ALLOWANCE

TOTALSATELLITE SYST

BRA-ON OTAssIUY CESSIUI I106 KG, IlDg KG, 1106 KG,

2.217

203

070

2.Sw

4.9m

1.039

1.262

0.850

9.794

,249

.rm

32.993

10.1m

43.152

131.1781

2.217 .2Lm

I.200

,230

14 loo

1.039

1.262

10.130

9.794

,240

.lGo

41.312

12.565

53.878

(22 559,

2.139

,203

1.109

,230

5.650

1.039

1.262

10.130

9.794

,240

.I00

32.693

10.057

42.780

GROW-W x

25.0

33.0

24.4

30.0

30.0

50.0

la,

30.0

23.1

75.0

0

30.0

1-i

Table 1.2-J. Nuclear Reactor Concept Weight Summary (Nov. 1977)

PRIMARY STRUCTURE

SEC. STRUCT.

ATTITUDE CONTROL SHEILDING

AEACTORS ,201

FUELPROCESSING

TURBO EGUIPMENT GENERATORS RADIATORS

POWER CONDIT. WIRE HARDNESS

ANTENNA

IMS EOUIP.

IMS CABLING ITS

PROPELLANT/YEAR SUBTOTAL

GROWTH ALLOWANCE TOTAL SATELLITE SYST.

II@ KG,

0.351

!.I12

0.20 0.54

2.05

1.91 Z0.W

3.34

1.83 11.94 I

1.839

0.60

9.00 0.081

0179

010 35 052 10.411

iiziz

x GROWTH

25.0

250

30.0

30.0 300

300

x)0

Jo.0

33.0 50.0

1M

23.1

500

loo

0

zB.B

CONCEPT WEIGHT COMPARISONS

POWER BASE “ch&VN’;;fN WEIGHT GROWTH ;:::T

1106 KG, 1x1 It@ KG1

CR-1 28.513 30.7 37.275

2 25.599 30.0 33.714

5 2slW 31.2 57.379

RANKINE 20.3Ei CS/STEAY

31.2 34.005

NUCLEAR 35.0% 29.0 45.15

‘G&EOUS CORE REACTORIMHD COULD POTENTIALLY REOUCE THIS TO 1.99 X 106 KG - REFERENCE: BTH IECEC PAPER 759010. ,973

AA0 14.95 X 1 d

ATOR OPTIMIZATION COVLO POTENTIALLY REDUCE THIS TO KG. CONOENOING STEAM RADIATOR (LOWER TEMPERATURE,

Rectenna. The ground receiver or rectenna transforms the transmitted radio frequency energy to dc current for distribution into the utility network. The area covered by a 5-gigawatt (GW) delivered power rectenna rate is shown for a typical city (Figure 1.2-3). The rectenna formed an ellipse that is approximately 6x10 km. An additional 4 km in radial length was provided to the security fence to assure a safe level of radiation outside the fence.

FirstCandidate Selection (April 1977)

The two concepts selected for further evaluation and definition at the end of the initial study in April 1978 were a photovoltaic (CR-2) approach and a variation of the proposed Rockwell Solar Thermal satellite. A summary description of the two selected point designs are given in the following two paragraphs. Both these concepts are described in greater detail in Volume II ,of the Final Report (SD 79-AR-0023, dated April 1978).

In addition the selected ground receiving station point design which differs slightly from the previous concept is briefly described below and in more expanded form in Volume II of this report.

l-9

Solar Photovoltaic (CR-2). The GaAlAs photovoltaic point design satellite power system concept is shown in Figure 1.2-8. The system utilizes aluminized reflector membranes to concentrate the solar energy on the cells. The satellite solar reflectors produce a concentration ratio of CR-2. The satellite employes the "Vee trough" configuration has three troughs per wing. The system has an overall efficiency of 6.08% and delivers 5 GW of electrical energy to the utility company on the ground. The overall planform dimensions are 3.85x21.3 km, and the depth is 1.69 km. The satellite has a mass of 36.56~10~ kg which includes a 30% growth factor for the mass. The system requires 30.6~10~ m2 of solar cells and 61.2~10~ m2 of reflector surface. The solar cells for the point design are GaAlAs cells rated at 20% efficiency at AM0 and 28'C. The solar array blanket mass is 0.2525 kg/m'.

Figure 1.2-8. Solar Photovoltaic Satellite (CR-Z) 5 GW (Apr. 1978)

Solar Thermal - Rankine. Figure 1.2-9 shows the Rockwell point design concept of a 5-GW solar thermal SPS using a cesium/steam Rankine cycle. The two concentrators are of an inflatable design, using aluminized plastic film with a transparent canopy. Sunlight is concentrated on an open-disc absorber (cesium boiler) which provides cesium vapor at 1260°C (2300'F) to cesium turbines. Exhaust from the cesium turbines is condensed at 593°C (llOO°F) on the outside of steam boiler tubes which produce steam at 538'C (1000'F) and 16.6 kN/m' (2400 psia) to a bottoming steam turbine. Exhaust from the steam turbines is condensed at 204°C (400'F) in a tube/heat pipe/fin radiator.

Each of the concentrator modules is hinged to permit seasonal tracking of the sun without imposing gravity gradient torques on the satellite. The beam connecting the two modules is offset to locate the rotary joint at the satellite center of gravity. This permits mounting of thrusters on the rotary joint and facilitates their orientation during LEO/GE0 orbital transfer.

l-10

. INFUTAILE

l x)0 PSI REFLECTOR FILM STRESS

l 5 KM DIA

*&sonBcR b 159 INDEPENDEI - POWER MOL”...

Figure 1.2-9. Solar Thermal - Rankine (5 GWJ (Apr. 1978)

Rectenna. The rectenna concept selected for further definition is illustrated in Figure 1.2-10. The receiving antenna forms an eclipse with major and minor axis of 13 km and 10 km respectively. The major axis is aligned along the N-S geographic line. Figure 1.2-11 illustrates the general site concept recommended by the study to date.

IECTLHNA SWFORT CONCEPT

RECTENNA MODVLE

Figure 1.2-10. Microwave Transmission Subsystem - Rectenna (Apr. 1978)

l-11

/sw’rcn IARD N0. ’ /CONVERTER L RELAY 1DG

-MAINTENANCE AtEA

1 CONTROL CENlEt (2‘03 d,

CONQETE RANT3

’ KtIMElEt fENCE

. RECIENNA IANCL tiEA ,I0 KY X I3 KM,

1 RECTENNA fMM ACCESS ROL9

‘SWITOI YARD NO. 3

Figure 1.2-11. Rectenna Site Concept (Apr. 1978)

Mass Properties. Table 1.2-4 and 1.2-5 present a summary of the estimated weight for the two point design concepts.

Table 1.2-4. Photovoltaic (CR-2) Satellite Mass Statement - Point Design (Apr. 1978)

Subsystem Keight

(Yfllion kg)

collector array

Structure and orchanisms Power source Power distribution and control Attitude control Infcrmacion nanagment and control

Total array (dry)

Anrenna section

Structure and mechanisms Thenal control Yicrowave power Povrr distribution and control Informtion management and control

Total antenna section (dry)

Tocal SPS dry weight

Crovth (30%)

Total SPS dry weight with growth

Propellant per year

3.777 8.831 1.166 0.095 0.050

(13.919)

1.685 l.LO8 7.012 3.:69 0.630

(11..?OL)

28.123

8.:37

36.560

O.OLO

l-12

Table 1.2-5. Solar Thermal Satellite Mass Statement - Point Design (Apr. 1978)

SUBSISTER

tOLLECTOR ARRAY

STRUCTURE AND HECHANISMS POWER SOURCE POWER DlSTRl8UTlON AND CONTROL ATTITUDE CONTROL THERMAL CONTROL INFORIlAT10N MANAGEMENT AND CONTROL

TOTAL ARRAY (DRY)

ANTENNA SECTION

STRUCTURE AND HECHANISHS THERMAL CONTROL HICROUAVE POWER POWER DISTRIBUTION AND CONTROL INFOIWATION MANAGEHENT AND CONTROL

TOTAL ANTENNA SECTION

TOTAL SPS DRY WEIGHT

GROWTH (30%)

TOTAL SPS DRY WEIGHT WITH GROWTH

PROPELLANT PER YEAR

WEIGHT (MILLION KG)

I.661 3.120 4.304 0.095 0.706 0.050

(18.016)

:% 7:012 3.469 0.630

(14.204)

32.220

9.666

(41.886)

0.040

NASA Reference Satellite Concept

In October 1978, NASA established a baseline (Reference) concept to be used in subsequent design and feasibility analysis. The primary approach selected consisted of solar blankets installed on a multi-trough, plannar structure with a microwave transmission system for power transfer to Earth located sites. The initial concept proposed a primary solar conversion approach utilizing Silicon solar cells with a concentration ratio of one (CR-l) and an alternate approach using GaAlAs with a concentration ratio of two (CR-2).

The Silicon cell based concept consisted of 8 cell troughs each containing 16 bays. The GaAlAs concept consisted of 5 troughs by 20 bays. Both concepts utilized an end mounted, 1 km (nominal) microwave antenna. Both concepts were normally 5.3x10.4 km, with the Silicon concept containing a greater mass. (51~10~ kg) compared with GaAlAs (34 10 kg). Figure 1.2-12 illustrates the GaAlAs version of the reference satellite. Overall system efficiency for the Silicon based concept is estimated to be 7.06%, while for GaAlAs the efficiency is estimated to be 6.97%. Power output for these concepts (at utility interface) is estimated at 5.0 GW.

Mass Properties. Table 1.2-6 presents a summary of the estimated mass for the two reference concepts.

l-13

i ,050 m

5

Figure 1.2-12. NASA Reference Configuration (Oct. 1978)

Table 1.2-6. NASA Reference Satellite Mass Properties (Oct. 1978)

SUBSYSTEM GAALAS CR = 2 SILICON CR = 1 OPTION OPTION

SOLAR ARRAY

PRIMARY STRUCTURE SECONDARY STRUCTURE SOLAR BLANEETS CONCENTRATORS POWER DISTRIBUTION & CONDITIONING INFORMATION MANAGEMENT k CONTROL

ANTENNA

PRIMARY STRUCTURE SECONDARY STRUCTURE TRANSMITTER SUBARRAYS POWER DISTRIBUTION IL CONDITIONING THERMAL CONTROL INFORMATION MANAGEMENT & CONTROL ATTITUDE CONTROL

ARRAY/ANTENNA INTERFACES*

PRIMARY STRUCTURE SECONDARY STRUCTURE MECHANiSMS POWER DISTRIBUTION

SUBTOTAL CONTINGENCY (25%)

TOTAL

l ROTARY JOINT. SLIP RINGS. ANTENNA YOKE

13.796 27.258

4.172 3.388 0.581 0.436 6.696 22.051 0.955 - 1.144 1.134 0.050 0.050

13.382 13.382

0.250 0.250 0.766 0.786 7.178 7.178 2.189 2.169 2.222 2.222 0.630 0.630 0.126 0.126

0.147 0.147

0.094 0.094 0.003 0.003 0.033 0.033 0.017 0.017

27.327 40.787 6.832 10.197

34.159 50.984

Recommended Alternative Satellite Concept

The Rockwell satellite concept as of December 1978 is presented in Figure 1.2-13. Figure 1.2-13(a) illustrates the Rockwell end mounted antenna while Figure 1.2-13(b) depicts a satellite with a center mounted antenna concept. Both approaches consist of a 3 bay by 10 bay structure containing the solar arrays and reflectors. The 30 bay structure is sized to dimensions of 3900 kg by 16000 meters. The center, antenna mounting, structure adds 1900 meters to the overall length of the satellite. The end mounted antenna concepts dry mass is greater by approximately 1.0X106 kg.

1-14

h : \

Cb)

Figure 1.2-13. Alternate Satellite Concepts (Dec. 1978)

The solar array panel is 600 m wide X 750 m long. Two of these panels make up a voltage string (45.7 kV). The 600 m width consists of 24 rolls each 25 meters wide. Sizing of the array is based on a solar constant at summer solstice (1319.5 W/m2), an end of life concentration ratio of 1.83, an operating temperature of 113°C and the design factors listed in the figure. A design margin factor of 0.975 is used to match the available area of 27~10~ m2. Total power at the array output is 9.52 GW. Total transmitted power is 6.79 GW. System efficiency factors for the satellite as indicated in Figure 1.2-14.

Mass Properties. Table 1.2-7 presents a summary of the mass for the two alternate concepts.

Ground Receiving Station. The various elements of the initially defined Ground Receiving Station (GRS) are shown in Figure 1.2-15. The major elements shown include the basic receiving/rectifying panels (rectenna), the power distribution and power conversion elements, as well as the various supporting elements (maintenance, facilities, land, etc.). The estimated efficiency of the various links of the ground system power chain is shown in Figure 1.2-14.

The rectenna panels are located in the center of the receiving station and covers a ground area of approximately 80 km2 (approximately 25,000 acres). An additional 32 km2 (approximately 10,000 acres) is required for the distribution and conversion stations plus a security perimeter. Received power is approximately 5.53 GW (at 2.45 GHz). Power available at the utility interface is approximately 4.6 GW ac.

1-15

(9.52 GW r ,

SWITCH SLIP

GEAR RING/ - RRUSH (~(8.92 Gw)

I- 0.9381- I POWER DIST

l- 0.7608 MW ANT

4

I \‘.-, ‘.,

I

GROUND

= POWER GEN, X POWER DIST. X MW ANT. X GROUND

(13.35%) (93.81%) (76.06%) (67.91%) b.47% (5.6% BASED ON TOTAL INTERCEPTED AREA)

I t7 SG = .w9’4 = .906

NOTE .W514 = .932 I

Figure 1.2-14. System Efficiency Chain

Table 1.2-7. Mass Properties - Alternate Concepts

OUTER

CONS1 ROAD

SUBSYSTEM END-MOUNTED

SOLAR ARRAY

PRIMARY STRUCTURE SECONDARY STRUCTURE SOLAR BLANKETS CONCENTRATORS POWER DISTRIBUTION G CONDITIONING INFORHATION HANACEHENT C CONTROL ATTITUDE CONTROL

(11.884)

.702

.55a 6.818 I.037 2.603 0.050 0. Il.6

l-037 0.882 0.050 0.116

ANTENNA

PRIMARY STRUCTURE SECONDARY STRUCTURE TRANSHllTER SUBARRAYS POWER DISTRIBUTION & CONDITIONING THERMAL CONTROL INFORMATION HANACEMENT & CONTROL

SUBTOTAL

CONTINGENCY (254)

(14.532) 0.120 0.857 7.012 4.505 1.408 0.630

(14.532) 0.120 0.857 7.012

26.416 24.557 6.604 6.137

TOTAL 33.020 30.694 I

500 KVAC BUS (TYP .) (3-PHASE,

FENCE\~~~~ \‘MONlTOR b. CONTROL FACILITY

iRUCllON ACCESS - . . A Y- NOT TO SCALE (PANEL AREA 10 KMX 13 KM)

Figure 1.2-15. Ground Receiving Station

l-17

1.3 TRANSPORTATION SYSTEM

Figure 1.3-l illustrates the baseline transportation flight operations designed to deliver cargo and personnel to geosynchronous (GEO) orbit for SPS construction. Three SPS unique elements of the system are: the Heavy Lift Launch Vehicle (HLLV), the Electric Orbit Transfer Vehicle (EOTV), and the Personnel Orbit Transfer Vehicle (POTV). The HLLV is a two stage parallel burn launch vehicle utilizing LOX/RP in the 1st stage and LOX/LH;! in the second stage. Second stage propellants are crossfed from the 1st stage during 1st stage burn. These stages take off from a vertical position and land horizontally in a manner similar to that of the Shuttle transportation system. Each HLLV launch can transport a 0.227~10~ kg (0.500~10~ lb) payload to low earth orbit (LEO).

A second major transportation element is the LEO-to-GE0 cargo transfer vehicle, the EOTV. The EOTV consists of a basic solar array structure and electric (ion) thruster arrays by which as much as 5.2~10~ kg of cargo can be transferred to a GE0 - located construction site. A maximum EOTV load would therefore require 23 HLLV missions.

A third vehicle is designed to transport personnel from the LEO staging area to and from the GE0 site. The vehicle consists of a single chemical propulsion stage and a separable crew module. The propulsion element is refueled in GE0 for return to LEO. Acceperation and operation restrictions are similar to those imposed for manned space vehicles.

Figure 1.3-l. SPS Transportation System - LEO Operations Operational Program

l-18

I -

1.4 PROGXAM GROUND RULES

Table 1.4-l shows the program ground rules that affected the development of requirements. Table 1.4-2 shows the general requirements describing the overall SPS program.

Table 1.4-l. Program Ground Rules

IOC DATE: 2000

PROGRAM SIZE: 2030-300 GU (IO CWYR)

SYSTEM LIFE: 30 YEARS

COSTS : 1977 CONSTANT DOLLARS (7.5% DISCOUNT RATE)

TECHNOLOGY BASE: 1990

SYSTEMS AVAILABLE IN THE 1980’S: SHUTTLE. IUS. 6 OTV

Table 1.4-2. General Requirements Describing Overall SPS Program

Progr-tic Technology -

ENERGY SOURCE-Solar

CAPACITY-Assume 2 units/year after initial buildup to 300 GW

LIFETIHE- 30 years with minimum planned maintenance (should be capable of extended life beyond 30 years with replacement)

rot-2000

OUTPUT POWER-Power level is defined as constant power level (except during solar eclipse) at utility interface (5 GW, nominal )

HAXIMLV4 RADIATION LEVELS--Maximum radiation level at rectenna is 23 mW/cm2; maximum radiation level at perimeter fence line is 1 mu/cm2 .

WEIGHT GROWTH-TBD BUILDUP-Provide 10 GW (nominal)/year power buildup rate to utility interface TOTAL WEIGHT-All sumnary weight (totals)

will be in t&n of kg/kw, OPERATIONS-Geosynchronous orbit; O-degree inclination, circular (35,786~km altitude)

RESOURCES-Minimum use of critical resources

COMWERCIALIZATION-Compatiblc with U.S. utility networks

DEVELOPMENT-Evolutionary, with provisions for incorporating later technology

ENERGY STORAGE-To support on-board satellite system operations only

CONSTELLATION-Satellite space, 3 degrees

FAILURE CikXTERIA--No single point failure may cause total loss of SPS function

STORAGE-One year on-board storage vithout resupply

CONSTRUCTION-Structural material to be graphite composite.

STARruP/.SHClTDCi#N-!fSD --

1-19

II

2.0 FUNCTIONAL FLOW BLOCK DIAGRAMS

2.0 FUNCTIONAL FLOW BLOCK DIAGRAMS

2.1 SATELLITE

2.1.1 INTRODUCTION

The functional flow diagrams as presented in this section are intended to provide two types of information for the satellite. The first is a simplified block diagram identifying the basic flow path of the end product, in this case electrical energy,from the point of origin (power generation) to the interface where the energy leaves the satellite (microwave antenna). The function flow diagrams of the ground receiving station are considered in section 2.2. The flow diagrams also identify the interfaces between the various primary and support subsystem and the signal flow paths within the various subsystem concepts. The flows are, at the present time, limited to the active paths necessary for vehicle operations. Passive elements or subsystems, e.g., the structural fasteners and fastening concepts, are not addressed in this section of the requirements document.

The second type of information provided is a summary of the requirements imposed by, imposed upon, and/or derived within the various subsystem configurations.

Examples of the type of information presented (when available) include operating temperatures, pressures, flow rates, voltage and current, pointing limits, etc.

The operating relationship between the various subsystems is illustrated by the block diagram shown in Figure 2.1-1. Operational control of the satellite is provided by the Information Management and Control Subsystem (IMCS).

The IMCS also provides subsystem processing support for all but the very special functions. The only specific case of a special function identified at the present time is the beam programmer element in the microwave antenna subsystem. The man-machine interface has also been established to be at computer-generated display/control terminals. The display/control terminals have not, as yet, been defined -- nor will they be during this contract.

Mechanical interfaces (between structure subsystem and the other subsystems) have not been shown to simplify the diagrams.

2.1.2 SUBSYSTEM IDENTIFICATION

The following paragraphs briefly describe details of each of the subsystem concepts defined to date. Important parameters are summarized where appropriate. More specific details of each subsystem element is discussed in greater length in Section 3.0 of this document.

2-l

DISRIDUlION

INFORMATION MLNAGEMNT AND CON,‘“01

Figure 2.1-1. SPS Satellite Subsystem Functional Relationships

Power Generation

Figure 2.1-2 presents the basic power generation concepts for the photovoltaic concept with a concentration ratio of 2 (CR-2). Switchgear for inter- segment connections are considered part of the power generation group.

TIE BUS

SLIP RINGS

42.95 K” 12.9 I(”

8,936 Gw t1.925 GW

; : SUMMING& BUS

21120 43.1 I(”

022 8.96 GW I Jv=emv

Figure 2.1-Z. Power Generation - Photovoltaic (CR-2)

2-2

Power Distribution

Figure 2.1-3 presents the functional block diagrams for the satellite power distribution subsystem. The supplementary power source is required during eclipse periods to power critical support systems and to sustain temperature-sensitive subsystem elements such as the MW antenna klystrons.

Figure 2.1-J. Power Distribution

Preliminary power estimates for supplementary power indicates a need for a l-40 MW/h storage capacity. Although subsystem power conditioning and dc-dc conversion is shown as being combined into a single unit, these functions are in actuality composed of many dc-dc converters (or dc-ac converters, if necessary) located throughout the vehicle and/or MN antenna structure. Dc-dc bias voltage converters are located at two locations on the antenna structure and supply the 5 high voltages needed to operate the Klystrons. The maximum number of high voltage dc-dc converters on the antenna is estimated to be 32.

Figures 2.1-4 through 2.1-6 present the subsidiary systems that make up the attitude control and stationkeeping subsystem for the photovoltaic SPS satellite concepts. These three systems are the attitude reference (platform) system, the microwave antenna pointing system (ring drive and gimbal drives), and the tank and engine systems,

2-3

r “1 r I MGMT I I -1

L-- J L-J a-

---a

1 AH CON? 1

pENG & 1Ky

I *RINGS L

GIW I

kINFIG 1

Attitude Reference System Figure 2.1-4. ACS - Photovoltaic

?? P

r--1 I**-1 WUIO IcoFllux I L w-d

r ff

Iwo 1 r--1 1 YGNT 1 I 1-I I

S-’ I

LW-.l -- L-w J

Figure 2.1-5. ACS - Photovoltaic and Solar Thermal, MW Antenna Pointing System

I- - - -1 lA~CONlROLt 1 (*nm \ I --,-I

7-7 c INK) 1 I MGMT I

r --1 r-- 1

Figure 2.1-6. ACS - Photovoltaic Tank and Engine System

2-4

.

Structure

Figure 2.1-7 presents the only active (instrumentation) portion of the structures subsystem defined to date. The depicted system monitors the location of corner reflectors so as to establish the degree of distortion existing in reflectors, mirrors, and other elements of the entire satellite configuration. Each of the 35 laser transits is assumed capable of scanning and calculating the location of at least 100 to 200 reflectors distributed over the surface of the mirror, solar arrays, and primary support structure.

13 r INFO ‘I 1 MGNT 1

I MO i LWNTROLJ --

--

LUUCIW~ I I L--J r-- 1

‘- I

I I L--d

Figure 2.1-7. Structure - Configuration Monitor - Photovoltaic (CR-Z)

Thermal

Figure 2.1-8 presents a general summary of the thermal requirements. As more information regarding temperature, status, etc., is defined, it will be added.

Microwave Antenna

Figure 2.1-9 presents the beam generation and control portion of the microwave antenna subsystem. Most of the paths shown operate at frequencies of approximately 2.45 GHz and are, therefore, either coaxial cable, strip lines, or waveguides. The beam programmer is a special-purpose dedicated processor design to accomplish high-speed RF pointing control via the digital diode phase shifter. External processing is limited to much slower, large element antenna .' pointing and performance monitoring and control.

Information Management and Control

This subsystem provides for overall satellite operational control, as well as performing system status monitoring.

2-5

- - ‘I

PHOTOVO'LTAIC, GaAlAs, CR = 2

*SOLAR ARRAY

I TBLANKET = ll3"Cl

*POWER DISTRIBUTION

*ROTARY JOINT

*INFORMATION MGMT

T MAX

= 60"~

El T

MIN = 60°C

*ANTENNA

CAVITY RADIATOR T = 200°C

COLLECTOR RADIATOR T = 700°C

Figure 2.1-8. Thermal Requirements

Figure 2.1-10 depicts the overall processor hierarchy appropriate to the basic photovoltaic (CR-2) configuration.

Figure 2.1-11 presents the typical architecture of the microwave antenna IMCS system.

2-6

I --

I- 1

I

I ro*n I , wanrtu, L- J

I T 30 wntnnw

Figure 2.1-9. Microwave Antenna - Beam Generation and Control

Figure 2.1-10. SPS IMCS Top-Level Block Diagram

N I

W

WaOvAVt

ANllNNA

COMNW I

RoJ

I

I-1 P _ rr I 7s

I MIQOWAVE

1 I

MICROWAvl

ANTENNA ___- ANTENNA I

RAC - RWOIF ACOUISI?ION AND CONlROL lJNll

SM - suR-hulllPlExER

)LP - buao4nocwoe

,a, - M. U)NlROL UN11

Figure 2.1-11. IMCS Microwave Antenna

2.2 GROUND RECEIVING STATION

2.2.1 INTRODUCTION

The functional flow diagrams for the Ground Receiving Station (GRS) are in many respects similar to those established for the SPS satellite. This becomes apparent when considering the relationship of the various subsystems and interfaces as shown in Figure 2.2-l. As on the satellite the Information Management and Control Subsystem (IMCS) provides primary control and system monitoring with the man-machine interface primarily used for judgement and system reconfiguration.

I , INFORhWION MANAGEMENT AND CONTROL

Figure 2.2-l. SPS Ground Receiving Station Subsystem Functional Relationships

2.2.2 SUBSYSTEM IDENTIFICATION

The following paragraphs briefly describe details of the subsystem concepts considered to date. Important parametric data are estimated and summarized where appropriate. More specific details of each subsystem element is discussed in greater detail in Section 3.0 of this document.

Power Reception

Figure 2.2-2 presents the basic microwave receiving/rectifying element (rectenna) located at the ground site. The receiver/rectifying elements (dipoles and rectifying diodes) are symetrically located on the panels as shown in Figure 2.2-3. Individual diode rectifier outputs are series connected to produce voltage strings slightly greater than 40 kV.

2-10

‘0.3lM WIDE FOOTING. 0.194 Ay3vL GUDL, 0.43 Y YLCW GRMIE (2 FUCES,

Figure 2.2-2. Basic Rectenna Panel Assembly

-7 A/2 = 0.428 M -)

t

-A----_ em-- -

SQUARE CLUSTER 49 DIPOLES --

OUTER PERIMETER 7A/2 -’

19.? BEAM AT 3 DB POINT AT 2% POINT

I f------

(ii& LOOKED AT HEXAGON, LINE)

INTERCONNECT SCHEME

2 LINKS 2 XgLONG

4 LINES 7/2 k 4 LINES 7/2 fiA

ALL OTHERS A g LONG

AVERAGE PATH = 4.2 A I TO DIODE

I ADDED LOSS L------- -_

.079 Dn 1.8%

DISTANCE - AVERAGE ELEMENT TO OUTPUT = 3/2 A8

. LOSS AT .015 Da/ A, = .0225 DB = 0.5%.

l POTENTIAL DIFFICULTY - MATCHING LOSSES OF 17 PORT NETWORK

Figure 2.2-3. Panel Dipole/Diode Cluster Layout

Power Distribution

Figure 2.2-4 presents the functional block diagram for the GRS. Power to supply the various operating systems during periods when the satellite source is not transmitting power, or during startup periods, is provided by cross- feeds to other auxiliary power sources not shown in this diagram.

2-11

96 DOUBLE BOXES EA SECTOR 1M . CUBES

\ 40,000 VDC BUSSES

I ‘-

14.8 MW

2 -g-c Eli IL

-----

ISOLATION SWITCHES \

I ( CONVERTER

I STATION I

w------------------

FEEDER POWER (INPUT TO 40,000 VDC BUSSES)

RANGE 12.1 - 16 MW AT

\

\

\

400 MW 500 KVAC

(lYPlCAL OF 12 SITES)

108s ROWS

(300 - 400A) ’ AS MANY FEEDERS AS REQUIRED TO APPROACH 16 MW

Figure 2.2-4. Ground Receiving Station Schematic Block Diagram - Preliminary

lr

Included in the power distribution networks are the voltage feeders run behind each rectenna panel, the 40 kV dc and 500 kV ac buses as well as the voltage string isolating motor switches and system protecting switchgear.

Power Conversion Stations

Figure 2.2-5 presents a simplified block diagram of solid state power conversion stations situated around the perimeter of the rectenna area. Initial power estimates result in a preliminary count of 12 stations for the 4.61 GW capability of a basic GPS site.

40 KVDC F,LTERS 40 KVDC N KVAC N KVAC TRANSFORMERS 500 KVAC

POWER F I LTERS - INVERTERS

z

4g

4 4 I I I

- I

- CONTROL _ _ - - - - - - - -I

Figure 2.2-5. Power Conversion Station Functional Block Diagram - Simplified

Structure

The basic structure of the rectenna panel is shown in Figure 2.2-2. Figure 2.2-6 illustrates a typical area of the rectenna farm. Also shown in Figure 2.2-6 is a panel installation mechanism that can be used during initial field buildup or to replace defec,tive or damaged panels during maintenance procedures. Details of support facilities, storage areas, and other required structures have not been established, and will be determined as part of other, yet to be established, studies.

Thermal

The specifics of thermal control and/or shielding have not, as yet, been determined. Details will be established by future studies (TBD).

Safety and Security

Elements of security and safety are TBD.

2-13

Figure 2.2-6. Panel Installation Operations

Information Management and Control

This subsystem provides for overall ground site control as well as all on site system status monitoring. With the addition of appropriate connnunica- tion channels the on site IMCS can also provide for off-site safety and security.

Figure 2.2-7 depicts a possible overall processor hierarchy appropriate to the needs of the GRS. Note that the selected architecture is similar to that selected for the satellite. This was done because the basic requirements (e.g., many parallel operations using relatively simple algorithms and very large numbers of measurements and controls), are similar.

The basic architecture of the individual subsystems have not been defined because of the limited details available.

2-14

Group A Equipment

I l Group C , Equipment

Group B Equipment

SUPERVISORY PUR CONY STA

SUPERVISORY SUPERVISORY

I

PERIMETER HON ITOR NO. 1 COMPUTER

I I

I I

I

I I

POWER DISTRIB POWER OISTRIB.

THERML GROUP A 4 CDHPUTER l

BCU

POWER DISTRIB. I

POWER DISTRIB. I

I THERHAL GROUP C I

Figure 2.2-7. IMCS Processor Hierarchy - Typical Ground

3.0 SUBSYSTEM

3 .O SUBSYSTEM

3.1 SATELLITE

The following subsections of this document describe the requirements, the major assemblies, characteristics, and definitions for the seven satellite subsystem groups which were developed during the continuing SPS evaluation. These subsystems (or subsystem groups) are listed below.

l Power Conversion

l Microwave Transmission

l Power Distribution and Control

9 Structure

l Attitude Control and Stationkeeping

l Thermal Control

l Information and Management Control

More detailed discussions/descriptions of the identified subsystems may be found in Volume II.

3.1.1 POWER CONVERSION

The baseline power conversion subsystem consists of solar cells, blankets, attachment devices, reflector membranes, and associated attachment devices. Gallium aluminum arsenide (GaAlAs) cells have been selected as the baseline solar cell. The cell is fastened to a thin-film Kapton substrate with an FEP adhesive. The photovoltaic power conversion subsystem is designed for a nominal geometric concentration ratio of 2.

Functional Requirements and Block Diagrams

The functional requirements for the photovoltaic power subsystems are listed in Table 3.1-l. The system efficiency block diagram is shown in Figure 3.1-1. Shown in the figure are power levels, efficiencies, degradation factors and solar cell area requirements. A simplified integrated block diagram for the CR-2 concept is presented in Figure 3.1-2.

Major Assemblies

The major assemblies and components that are required for the photovoltaic subsystem are shown in Figure 3.1-3.

3-l

Table 3.1-l. Solar Array Functional Requirements

PROGRAMMA'TIC --- ENEIGY SOURCE - Solar CtlPACI'I'y - 9.5 GW (nominal) delivered to power distribution networks LIFETIME - 30 years with minimum planned maintenance (should be capable of extended life beyond 30

years with replacement) K-K' DA7'E - 2000 OPERd'TIONS - Geosynchronous orbit; O-degree inclination, circular (35,786 km altitude) RESOURCES - Minimum use of critical resources COPlEfERRCIAIJZdTION - Compatible with United States utility networks DEVEU3PEfENT - Evolutionary, with provi:;ions for incorporatinq later technoloqv

r -.. _-.- . __ TECHNOLOGY J

OUTPUT POWER - Power level is defined as constant power level (except during solar eclipse) MASS GROWTH - 25 percent ENEH(;Y STORAGE - To support on-board satellite system operations only FdIfJlRE CRITERIA - No single-point failure may cause total loss of SPS function ENERGY PAYBACK - Less than 3 years COST - Competitive with ground-based power generation within lifetime of SPS project STORAGE - One year on-board storage without resupply

Standby (zero power) Turn on after leaving eclipse and arrays reach equilibrium temp-

edundant operation, auto shutd

*Solar cell/blanket/reflector module

.999

-I - r=*$@ -I ,

SWITCH _ HI VOLTAGE SWITCH

RESONANT

GEAR DC-DC - - KLYSTRONS

SWITCH _ CAVITY +,

CONVERTER GEAR

RADIATOR 1

3 - .999 .96 .592 .85 .96 I 7.07 GW

b 6.79 GW 1

I- r= .98 -l!

4.61 GW’ I

SWITCH _ POWER SWITCH I r?i~TZ;iy I

GEAR CONVERSION GEAR - RECTENNA 1 A~~,&YEPE r] llL\L r= TRANSMISSION EFFICIENCY L,,-- J

.998 -96 .997 .89 ,815

Figure 3.1-l. System Efficiency Block Diagram

F O.l*FOR CC

SOLAR ENERGY 1 1

~RCIJI, DISPLAYS 1 Ifi

1

PHOTOVOLTAIC POWER CONVERSION GoAl*l CELLS ZWa *MO WC

SP.SS GW. h CONTROL

Figure 3.1-2. Simplified Integrated Block Diagram - Photovoltaic (CR-21

3-3

SOLAR PHOTOVOLTAIC ’ POWER CONVERSION

I I

SOLAR CELLS SOLAR BLANKET CONCENTRATOR

l &A$ CELL l KAPTON BLANKET l KAPTON BLANKET

’ GaAIAs WINDOW l TRANSPARENT THERMAL .* REFLECTIVE BLANKET’ COAT INC

l COYER/SUBSTRATE l ADHESIVE l Al-TACHHENiS AND TENSIONING DEVICES

l CURRENT COLLECTORS l INTERCONNECTS l SENSORS

l AIR COATING ’ ATTACHHENTS AN0 TENSIONING DEVICES

l SENSORS

Figure 3.1-3. Assembly Tree - Solar Photovoltaic Power Conversion

Solar Cells. The solar cell used in the SPS design is a GaAlAs cell having an efficiency of 20 percent at Air Mass Zero (AMO) and 28°C. The cell consists of the GaAlAs junction, GaAlAs window, cover/substrate, current collectors, and an anti-reflection coating. The basic cell design is the invert- ed GaAs/sapphire designhaving a weight of 0.252 kg/m'. The various cell designs and the selected design are shown in Figure 3.1-4. The design cell has a 2O+m sapphire substrate upon which is grown a 5-pm single crystal GaAs junction. A 500-g GaAlAs window is then deposited on the 5-urn junction. The voltage and current characteristics of the cell as a function of operating temperature are shown in Figure 3.1-5. The cell build up to form a submodule for the solar blanket is shown in Figure 3.1-6.

Solar Blanket. The solar blanket consists of a 25-urn Kapton membrane upon which the cells are fastened with a thermosetting FEP adhesive. Also included in the blanket are the interconnects, transparent thermal coating as may be required for thermal control, attachments, tensioning devices, and sensors. The solar cell blankets will be manufactured in blanket form and the solar cells attached. This assembly will then be rolled up on a drun type canister. It is postulated that the blankets will be 25 m wide by approximately 750 m in length. The canisters are then transported to orbit where the blankets are deployed via a roll-out deployment-type operation. The solar blanket consists of 1 square meter modules that are hooked together in series and parallel and the voltage, current, power output and a typical layout is shown in Figure 3.1-7.

3-4

-

Figure 3.1-6. Alternate Solar Cell Design

.46

50 loo WI MO 25Q

SOUR CELL lEMFE9AlLRE f-0

TODAYS XIMT CELLS (HUGHER S9S DESIGN S.O.L.

l2=20.4 (Z.W.J.59 04 CR-I; 12x 8X-1.72; I2SC

.W3.1% (Aho 47.6%

V~9-0.69 VOLlS/CELL V~-x69V/(ILLm.m2v/ch4l ly+27MA-U.5WCMWlDTM I,&23 M*JCM WIDTHI (10.5.’ MAfCUl

( b-f01

Figure 3.1-S. GaAlAs Solar Cell Voltage and Current Characteristics

3-s

INlERCONNfC75 &TO, GRID C.ONlACTS

‘25 PM KAPTON

\ 0lANKEl-l

6 b+POlYMER 1HERMAL COATING

rnlGH1 fMGm.4

7.w

a.4

0.03

1.66

4.0

1.7

3.4

o.(

25.25 t.252 KG/M+

Figure 3.1-6. GaAlAs Solar Cell Blanket Cross Section

II I27nsU II SOLAR ARRAY DESIGN FAClORS ---.-.. .._. -, ENERGY ONTO CELL (CR=1 .q ii1417

b , REFLECTOR,

OF!5 TEMP I l3C qrn 438.5

SOLAR CtLL DEkiGN FACTOR t.89 390.w SEASONAL FACTOR t.968) 377.70 DEGRAD FACTOR (.96) 362.67 SG FAClORt.997) 361.58 MARGINL975) 39.6

SOLAR ARRAY POWER OUTPUT - 352.6 W/M2 X 27.0 (lo6, M2 = 9.52 GW

Figure 3.1-7. Solar Panel Power Output - Watts/m2

3-6

Concentrators. Thin reflector membranes are used on the SPS to reflect the sun onto the solar cell surfaces and obtain a nominal concentration ratio of 2. The reflector is made of 12.5 pm (0.5 mil) aluminized Kapton. Reflec- tivity of the reflector was taken at 0.9 BOL and 0.72 EOL. The reflector membrane has a mass of 0.018 kg/m'. The reflective membranes are mounted on the structure using attachments and tensioning devices. Tensioning based on structural limit of the existing beam design (with safety factor of 1.5) indicates that tensioning of up to 75 psi can be used.

Design and Performance Characteristics

The design and performance characteristics of the photovoltaic system are presented in Tables 3.1-2 and 3.1-3. Operational parameters, materials of construction, deployed and planform areas and weights are presented for the subsystem.

Table 3.1-Z. GaAlAs Solar Cell and Blanket Preliminary Specification (CR-2)

ITEM CHARACTERISTIC

ARRAY INTERCEPTED ENERGY 69 GW CELL 5 AT 28’C, AH0 20% CELL r, AT Il3”C, AH0 18.15% ARRAY OUTPUT TO DISTRIB. BUS EOL 9.52 GW ARRAY OUTPUT VOLTAGE 45.7 kV CELL OLITPUT VOLTAGE AT Il3’C 0.7 v CELLS IN SERIES 65.000 SOLAR CELL SUEPANEL SIZE 600x750 m NUMBER OF BAYS PER SPS ARRAY DESIGN FACTOR & REFLECTIVITY & DEGRADATION 0.90 BOL. 0.72 EOL

CONCEh’TR,4TION RATIO GEOI’IETRIC 2 EOL 1.9 EOL 1.86

SOLAR CELL CONSTRUCTION COVER 20 urn SAPPHIRE CELL 5 urn GaAlAs INTERCONNECT 12.5 pm SILVER MESH SUBSTRATE

ADHESIVE 12.5 urn FEP FILM 25 urn KAPTON TRANSPARENT THERHAL COATING 6 urn POLYHER

SPECIFIC WEIGHT 0.2525 kg/m’(0.0516 lb/f?)

DEPLOYED CELL C BLANKET AREA PLANFORH 62.4 km2 SOLAR CELL AREA 27 kd REFLECTOR SURFACE AREA 54 km2

MASS SOLAR CELLS 6.818~10~ kg REFLECTORS 1 .037x106 kg

TOTAL PASS 7.855~10~ kg

3-7

Table 3.1-3. SPS Reflector Preliminary Specification (CR-L?)

ILeE

Material Kapton thickness Kapton specific gravicp Aluminized coating thickness Weight of aluminized coating

Reflector surface protective film coating

Reflector subpanel size

Number of reflector panels Reflector reflectivity/degradation Concentration ratio geometric Concentration ratio Reflactor slant angle from horiz.

Operating temperature Top reflectors Inboard bottom reflectors Outboard bottom reflectors

Total area of reflectors

Total weight of reflectors

Aluminized Kapton 12.5 m 1.42 (0.013 kg/m') 400 angsrrom uni:s 96 k&m'

Quartz or calcium r':u,>riie 600x 750 m

120 0.90 BOL 0.72 EOL . 2.0 1.9 BOL. 1.86 EOL 60 degrees

-52'C 46°C -73OC

61.2~10" m'

1.012~106 kg

Subsystem Definition and Interfaces

The subsystem interfaces are shown in Table 3.1-4 for the photovoltaic conversion subsystem. The major interfaces include the array orientation, attitude control, IMS and control, energy storage, power distribution, structure, thermal control, and support operations.

3.1.2 MICROWAVE POWER TRANSMISSION SYSTEM

The microwave power transmission system (MPTS) consists of a set of dc- microwave conversion devices, feeding a microwave array and a ground array of antenna/rectifier assemblies for microwave-dc conversion (rectenna). The ground array will be discussed in Section 3.2.

The array is phased by means of a pilot beam formed at the rectenna which is received by the array and used to form the power converter drive signals.


A functional block diagram of the satellite array assembly is shown in Figure 3.1-8. The requirements for the various operating modes are listed in Table 3.1-5.

Figure 3.1-9 shows how the array is formed of mechanical assemblies supported by two grids of catenaries anchored to the hexagon frame. These assemblies consist of nine 10.2 m by 11.64 m subarrays. Each subarray is formed of a variable number of power converter/radiator modules, depending on the power density required.

3-8

Table 3.1-4. Solar Array Interfaces I Subsystem Interface

r Interface Requirement Value or Comment

Voltage Power (E.OL)

45.i kL’ 9.52 Cli I Power distribution

+-- --

I -l-kermal

Array Reflector

Structure

Temperature Temoerature

Misorientation and mis- alignment angle

Orientation

Deflections

Control and monitoring of power subsystem

113°C 46°C to -73’C

0.10 oaximum Long axis perpendicular to orbital plane

Apply forlles to ade- quately tension array and reflector up to 75 psi Reflrccor deflections/ misalinnment cO.1'; array deflectionlmis- alignment <3"

TSD

DNA an

Figure 3.1-8. Functional Block Diagram

I t UmAlloN WCWNICY

counu

3-9

Table 3.1-5. Microwave Antenna - Operating Modes

asing tests Rectenna tests

a MECHANICAL

'GRAPHITE COMPOSITE COI@RESSION FRAME 'COMPOSITE TENSION WEB 021 Kwp2 RADIATION AT CENTER 050 KW PER KLYSTRON (136,000 KLYSTRONS)

Figure 3.1-9. Satellite Antenna Array Assembly

Major Assemblies

Figure 3.1-10 illustrates the major assemblies comprising the MPTS. Figure 3.1-11 shows a high-density module at the array center.

The selected power converters are nominally 50-kW klystrons, mounted in the center of the resonant cavity radiators (RCR). The klystron collector radiates both downward in the direction ?f the microwave rectenna as well as

i-10

. KLYSTRON (I) . WAVEGUIDES ( 1) ~-~?-~~<-~~~~ft~~ ~yij~&y--- ---; --w---w

(4) ~~~~TR-O-~----- l ELECTRONIC

-,,,-,,,-,J BEAM STEERING l SENSORS l REFERENCE

SIGNAL FEED

(2) r~-k7~Xi7 X5 ?%7 7 L- -- ---,,,,A l PHASE CONTROL l WAVE GUIDE PANELS

(3)~~r&Kc7o~Xoi~ -------a l RiFiRENCE SIGNAL (1) SEE POWER DISTR. d CONTROL

GENERATION (2) SEE THERMAL CONTROL l SENSORS (3) SEE ATT. CONTROL

(4) z ??%Tk’o’iS- - - - y --------,..a (4) SEE Ih’CS (5) SEE SXUCTURES

Figure 3.1-10. Assembly Tree - Microwave Power Subsystem

TYPE PWR DENSITY 1. 21.05 KW/M2

l NO. OF KLYSTRONS/ .5Q SUBARRAY

10x5

ARRANGEMENT

.PWd MOD SIZE 1.02 l HINGED

x CONFIG

2.33

.SHIPPING SIZE & WT. 10.2M x 2.33 2.35 x 1.02 x 2.33 7 6.8 KG

Figure 3.1-11. Klystron Subarray Assembly

to the rear of the radiator, as shown in heat from the klystron body and transfer between the radiating slots of the RCR.

Figure 3.1-12. Heat pipes remove it to the RCR face. The pipes lie

, HEAT PIPES

MICROWAVE SLOTS IN RCR

ELECTRONICS MOUNTING IS0

KLYSTRON COLLEClOR

Figure 3.1-12. Heat Radiators on Array Face

The transmitted signal is formed from the pilot beam by means of the retroelectronics shown previously in Figure 2.1-9; there is one of these circuits per subarray. Figure 3.1-13 shows a servo system for transferring the required reference phase from a central point to a mechanical module, where it is distributed to the nine subarrays.

f 4

RWRENCE SIGNAL 06lRINllON TRANMSSICW LINE

RUSE

stwrR

M

-3 DaJuR

b

Figure 3.1-13. Reference Phase Distribution System

3-12

Figure 3.1-14 illustrates the power supply system required for each 50-kW klystron. Note that the "mod anode" is a low-current electrode. It would be supplied by a separate circuit capable of varying its potential to control klystron power.

PtJ 0.5 w PO 50 KW

MOD-ANODE

POWER SUPPLY

(REGULATED)

P =3.2 KW 1.4-

ELECTRON BEAM 0.7’

POWER SUPPLY + 0.47

0.08 A

‘(REGULATED)

I COLLECTOR 1

/COLLECTOR

I POWER SUPPLY

OVERALL EFFICIENCY ? = 84.596

(NON-REGULATED) PT = 59 YW

Figure 3.1-14. Klystron Power Requirements (Preliminary)

Figure 3.1-15 is a layout and perspective view of the microwave integrated circuit (MIC) assembly which forms the transistor amplifier used for the alternate power conversion method. Figure 3.1-16 shows a possible circuit schematic of a solid state amplifier which consists of a puch-pull emitter follower driv- ing a common base puch-pull final amplifier. All four transistors are formed on a single chip as shown in Figure 3.1-17.


The functional requirements for the MPTS system are shown in Table 3.1-6 through 3.1-11. Table 3.1-6 summarizes the system functional requirements; Table 3.1-7 shows the prime power requirements for the array; Table 3.1-8 shows a phase error budget for the retroelectronics; Table 3.1-9 shows the array characteristics; Table 3.1-10 shows the characteristics of the klystron power module; and Table 3.1-11 shows the characteristics of the alternate transistor power module.

3-13

FIVE TO TEN _

TO SECOND DC CONNECTOR COMBINER OR ANTENNA

HOUSING WITH SAPPHIRE POWER DIVIDER/COMBINERS AT BOTTOM.

POWER MODULE CTD, ,rm IBC

POWER MODULE BASE TEMPERATURE: 275OC TO 285°C (GAUSSIAN) (PRELIMINARY cA~cuti~10~s) 160Y To 170% (UNIFORM-~~RGE~I ARRAY)

Figure 3.1-15. Transistor MIC Amplifier

Figure 3.1-16. Transistor Power Circuitry

3-14

STRIP LINE SLOTS

NOTE: DIVISION INTO CELLS NOT SHOWN

I , 1

Figure 3.1-17. Transistor Chip Layout

Table 3.1-5. Functiqnal Requirements

Power conversion efficiency (dc-RF) 0.85 Array radiation efficiency 0.96 Array beam efficiency 0.85 Atmosphere transmission efficiency 0.98 Rectenna conversion efficiency 0.88 Rectenna aperture efficiency 0.98 Total MTS efficiency 0.593 Harmonic level required, dB TBD AM noise decrease required, dB/kHz TBD PM noise decrease required, dB/kHz TBD Maximum power. density, kW/m' 21 Temperature at back of'array 60" Array mass TBD

DIPOLE

ED THRU)

I DIPOLE

1 3-15

Device

Table 3.1-7. Design and Performance Characteristics

Voltage Regulation Frea. W) Current A Power (kTJ) (73

Collector 1. 40 Collector 2 32 Collector 3 24 Collector 4 16 Collector 5 8 Klystron body 40.0

i Mod anode 20.0 ~ Cath. heater 20 v

0.28 0.35 0.47 0.70 1.40 0.08

11.2 11.2 11.2 11.2 11.2

3.2

0.1

None None None None None

10 10 10

Solenoid 20 0.5 1 Computer (1) 20 0.1 1 Retroelect. 20 v 0.1 1

Total 59.2 84.5

(1) Not included in determining klystron efficiency.

Table 3.1-8. Phase Error Budget

5OO:l power divider 6" Ref. 0 dist. link 6" Zone feeder 6" Retroelectronics 6" Klystron 0 shifter loop 3" Subarray pointing loop 3"

Total RMS phase error = 13"

Table 3.1-9. Array Characteristics

Operating frequency 2.45 GHz Operating wavelength 12.2 cm Mechanical module size 34.92x30.62 m Subarray size 11.64x10.2 m Subarray beamtiidth 0.73" Subarray RMS phase error 50-10" Amplitude weighting 10 dB Gaussian tape] Amplitude quantitization 10, 1-dB steps Subarray weighting Uniform Electronic steering limit, l/8 subarray beamwidth = 0.1" @MS) Mech. subarray pointing accuracy 0.07" @MS) Mech. module pointing accuracy 0.07O Total RMS subarray electronic pointing accuracy O.1° Polarization Linear Beam efficiency 0.85 Radiation efficiency (0.2 dB feed + 0.2 dB RCR R loss) 0.96 First sidelobe level -25 dB Electronic subarray steering accuracy,l/30 (0.73O) = .024"

(Pointing Loss = -.005 dB)

3-16

Table 3.1-10. Klystron Power Module

Output power, kW 50 Overall efficiency, % 85 Input power, W 0.5 Collector power dissipated, kW 5.0 Gain, dB 50 Body power dissipated, kW 3.3 Basic tube efficiency, % 85.8 Second harmonic level, dB -40 Prime power, kW 58.3 AM noise, dB/kHz -140 Auxiliary power, kW 0.70 PM noise, dB/kHz -130 Total power, kW 59

S~.~&ystem Def .nltion and Interfaces

Table 3.1-11. Transistor

Amp output power, W 120 Gain, dB (two stages) 20 Input power, W 1.2 Collector efficiency .83 Overall efficiency .78 Second harmonic level, dB -40 AM noise, dB/kHz >-140 PM noise, dB/kHz >-130 Number of amplifiers 100 Power module output, kW 12

Subsystem interfaces are shown in Figure 2.1-9. The concept/subsystem illustrated will be identical regardless of which satellite concept is selected for further consideration.

3.1.3 POWER DISTRIBUTION AND CONTROL

The power distribution and control subsystem (PDS) receives power from the power generation subsystem, and provides the regulation and switching required to deliver regulated power from distribution to the antenna system (Klystrons) and the various subsystems (Attitude Control, IMCS, etc.). During the ecliptic periods, batteries will be utilized to supply the minimum required power to the various subsystems. The feeders, and power cabling of all SPS subsys terns, are included in the PDS. The grounding, electromagnetic interference control, and shielding requirements of the SPS are also included as part of the PDS. The life expectancy of the PDS is 30 years with the exception of the energy storage (batteries), which has a life expectancy of 10 years. Resupply of the PDS will be as needed.

Fu~c~ttpnal~ Remquirements and Block Diagrams

Functional requirements for various operating modes are listed in Table 3.1-12. A simplified block diagram for the photovoltaic concept is presented in Figure 3.1-18.

3-17

Table 3.1-12. Power Distribution and Control-subsystem - Operating Modes

Mode

Construction Inter-orbit transportation

Operaticn EdiF=

Transition (from con- strwtion trms3ort.l L-Z _______ -- --

Failurcy ndstenancti - --- Cixtckouc

Major Assemblies

Assembly i .Function i

lied to subsvstems

Figure 3.1-19 illustrates the major assemblies comprising the power distribution and control subsystem (PDS).

Power Distribution. The power distribution subsystem consists of the main feeders, secondary feeders, summing buses, tie bars, and power interface cabling for the various subsystems. The main feeders are generally sized to minimize the combined mass of itself and the solar array mass, considering power requirements, efficiency, and the variation in resistivity with operating temperature. The power distribution system utilizes flat aluminum (6001-T6) feeders where feasible, and round conductors for those subsystems where flat conductors are not feasible. The flat conductors are not considered part of the main structure; they will normally be passively cooled by radiation to free space.

Regulation. The solar array output will be regulated so as to prevent line surges when switching the solar array power on to the main feeders. The regulation function is accomplished by selective control of intra-blanket switching managed by the information management and control subsystem (IMCS).

Power Converters and Conditioners. The power converter and conditioners convert the existing bus voltages to the subsystem voltage required for the various subsystem loads. The output tolerances will be based on the using subsystem interface requirements. The power converters are utilized in the GE0 mode of operation.

Switchgears. Switchgears are used for:

l Isolation of solar array blankets due to systematic element failure

l Isolation of solar array blankets when performing maintenance work

8 Prevention of large line transients upon startup and shutdown and during ecliptic periods

3-18

I SUBSYSTEM LOAD IUS I-

MAIN FEEDERS

SLIP RING “*”

TYPICAL .Q DRLISHES

TO SYSTEM “I” -TO SYSTEM ‘1” KLYSTI

DEVICE VOLTAGE

COLLECTOR 1 OKV

2 22KV

2 24KV 4 16 KV 6 8KV

KLYSTRON MDV 40KV YOD ANODE ?oKV

CATHODEHEATEI IOV

SOLENOID 2GV COMPUTEn 1oV

RETROELECT 2OV

TOTAL

SIA SOLAR ARAAY

k. SWITCH GEAn SOLID STATE CIIK IRKR.

iT REGULATOR IATTEnV TIE CONTACTOII

Is KlWEl

KW

11.2

11.2

11.2

11.2

11.2

a2

01

DL

01

01

60 8

EGULA. TION

1 II

1X

1m

ID6

10x

IOI

Figure 3.1-18. Power Distribution - Simplified Block Diagram

DlSTRlSUTlON

l MAIN FEECERS l SECOf;OARY

FEEDERS -- _-.-. l BUSES l SENSORS

POWER DlSTRlBUTlON & CDNTRDL SUESYSTE3

l BATTERY CHARGER . BATTERY

.* BOOST CONVERTZR

r- -I

. VOLTAGE l GEAR 9 SENSOR o CONVERSION l MONITllR

l RECTIFIERS . SOLID STATE l CIH. BREAKER 1 . SENSORS

. BRUSHES

. SLIP RINGS l SHOES . SE&SO RS

(1) -.- 1 I SECONOARY -

_ STRUCTURE I

(1) SEE STRUCT SUBSYSTEM (2) SEE INFO. MGMT AND

CONTROL (3) CONTROLLED BY IMCS (4) MAJOR INTERFACE W!TH

SOLAR PHOTOVOLTAIC POWER CONVERS I ON SUBSYSTEM

*INSTALLATION

Figure 3.1-19. Assembly Tree - Power Distribution and Control Subsystem

The switchgears will be solid-state to reduce the overall mass of switches. The voltages and currents being handled by these switches will be monitored by the INCS to determine their status and to establish a need for the opening and closing of these switches. The switches are generally held in the closed state during the steady-state mode of operation. During the startup and shutdown operations, the switches will be monitored by the IMCS and when certain voltage levels are reached a command signal will open or close switches as required.

Energy Storage. Batteries will be utilized during ecliptic periods to provide the minimum energy required by the various subsystems. The batteries will be a sodium chloride type, having a density of at least 200 Wh/kg.

3-20

Rotary Joint. The rotary joint is utilized to transfer energy through slip rings and brushes from the SPS fixed member to the SPS rotating member upon which the microwave antenna is located. The power transferred includes both that required to operate antenna-mounted equipment, as well as that to be transmitted to the ground.

Control. The PDS control concept is a simple, continuous monitoring system performed by the on-board IMCS computer system. The IMCS monitors the bus as well as the converter voltages, currents and temperatures, and compares these with preset levels stored in the computer(s). In the event of a voltage/current level disagreement with the preset conditions, the IMCS will initiate a cormnand signal to regulate the faulted area by opening up or closing the associated switchgear(

Secondary Structures. Secondary structures consist of mounting brackets, clamps, and installation structures as needed. It is assumed that a delta of 10 percent of the PDS mass would be sufficient for such purposes.

Design and Performance Characteristics _--- -~-~. .-~~

The design and performance characteristics for the power distribution subsystem are listed in-Table 3.1-13.


V~jor AsseTbly Requirements Technology Issu< -.---- --

GENERAL Mass Configuration dependent MTBF Subsystem dependent Life 30 years Efficizzy 88-98% (config. dependent) Resuppl:- and maintenance As needed

POWER DI~?.IBUTIO?I (PD) Mostly flat conductor Further study is Mass Configuration dependent required to deter- Materihl Aluminum 6001-T6 mine feasibility L>f Insuls:ion l-mm Kapton superconductivity Efficiczzy 88-98% (config. dependent) for reduction of Subsyszzx cabling Location and power dependent mass. ResuppI:- and maintenance As required Life 30 years or greater

POWER CO!.-:sITER .UTl COXDITIONING Dens icT 0.197 kg/kW Further analysis is Voltape Subsystem dependent required to specify Currecz Subsystem dependent design requirements Efficii-xy 96-98% and type. Life 30 years Resup?::: and nainienance As required

SWITCH GE1-Z Densic:; Approx. 0.00086 kg/kW Study is required ti 5w Penning discharge tube specify design Power raring Configuration dependent requiraments. Voltaj Config. and location dependent Efficiazcy 99-99.9% Life 30 years Resup;::; and maintenance As required

ENERGY SI:~GE (BATTERY) Densic~ Approx. 200 Wh/kg Further study to Type Sodium chloride define chargefdis- Temperature 2oo"c charge cycle, size, Efficiency 80-95% (turnaround) volume, and instal- Life lo-20 years lation is required. Resupllr and maintenance As required

3-21

Table 3.1-13. Design and Perfdmance Characteristics (Cont.)

SECOKDARY STRUCTLXE Mass 1OZ of PDS veig:lt was considered

equired for maunting and

Temperature sensors Current sensors Voltage sensors Switch gear control

Overcurrent Overvoltage Undercurrent Undervoltage

No. of sensors config. dependent No. of sensors config. dependent No. of sensors omfig. dependent Configuration d+endent

Slip rings required for SPS Operating voltage

2 positive a-C 2 negative

cross-section

d/each slip ring

Shoe size Dimension/shoe Contact surface area/shoe Weight/shoe Shoe travel velocity Wear rate per year Current density Operating teoperature

Sumyscem ueflnitlon ana mterraces

Subsystem interfaces are shown in Figure 2.1-2 for the photovoltaic concept. The power required from the photovoltaic power source is 9.55 GW.

3.1.4 STRUCTURES SUBSYSTEM

The primary SPS structure assemblies are made up, basically, of tribeam girders, tension cables, and joints. The fabrication and assembly of these structures are accomplished on orbit by beam machines and supporting auxiliary equipment. These structural elements must individually withstand the forces, torques, and dynamics imposed by the construction process. Once built up to an assembly level (e.g., solar array wing, rotary joint, etc.), the structure

3-22

must have sufficient strength and stiffness to withstand forces, torques, and dynamics generated by the environment (gravity-gradient torques), the attitude control system (forces and frequencies) and the operational equipment (rotary joint torques, microwave induced thermal environment, etc.). The level of strength and stiffness are dictated by other subsystem requirements such as pointing accuracies and ACS bandwidth frequencies.

The secondary structure consists of the passive interface attachment between the primary structure and the operational subsystems. The structural mechanisms consist of active structural subassemblies that articulate, rotate, or otherwise cause or allow motion between the primary structure and other subsystem elements or between subsystem elements themselves.


Functional requirements for various operating modes are listed in Table 3.1-14. Since the structure is primarily a passive system (the exception is the figure monitoring system), no block diagrams exist. A simplified interface diagram is presented in Section 2.1.2.

Table 3.1-14. Structural Subsystem - Operating Mode 1

MODE ASSEMBLY FUNCTIONS

MAINTAIN STRUCTURAL INTEGRITY OF STRUCTURAL CONSTRUCTION SYSTEH SUBELEMENTS PRIOR TO OVERALL SYSTEM STABIL-

IZATION.

PRE-INTER-ORBIT

I

WITHSTAND GRAVITY-GRADIENT/ACS TORQUE INTER-

TRANSFER REORIENTATION SYSTEII ACTION. DEVELOP f4lNlHUH F.lRST BENDING MODE FREQUENCY DICTATED BY ACS.

I I WITHSTAND G-LOADS ASSOCIATED WITH PROPULSION ---I

INTER-ORBIT TRANSFER I

SYSTEM I

SYSTEM THRUST LEVEL WHILE MAINTAINING ADEQuATE RIGIDITY WITHIN POINTING TOLERANCES REQUIRED I

I I FOR POWER CONVERSION. I

ECLIPSE I

WITHSTAND THEWAL TO EXTREME TEMPERATURE CHANCES.

ANTENNA STRUCTURE POINT UITHIN +0.08’ OF TARGET

OPERATION SOLAR ARRAY HAINTAIN WIDTH DISTORTION 2 ‘_0.3*

PROVIDE FOR SUBELEMENT SECONDARY LOAD PATHS. FAILURE/MAINTENANCE SYSTEM WITHSTAND FORCES C TCRQUES INTRODUCED BY

HAINTENANCE OPERATIONS.

Major Assemblies

Figure 3.1-20 depicts the major structural subsystem assemblies and tabu- lates the elements that make up each of these major assemblies. An example of this element breakdown in shown in Figure 3.1-21.

3-23

Figure 3.1-20. Assembly Tree - Structures Subsystem

*BASIC BEAM ELEMENT

Figure 3.1-21. Structure Breakdown

3-24

DesiEand Performance Characteristics -

The design and performance characteristics for the structures subsystem are listed in Table 3.1-15.


Subsystem interfaces are shown in Figure 2.1-7. The only active interface identified to date is the laser transit network, established to determine the satellite figure for the CR-2 photovoltaic satellite. It is expected that this network would be applicable to any photovoltaic concept.

3.1.5 ATTITUDE COXCROL AND STATIONREEPING SUBSYSTEM

The attitude control and stationkeeping subsystem (ACSS) is an integrated system designed to satisfy the contr.01 and stationkeeping requirements for each of the SPS operational modes. The functional performance requirements of the ACSS are to provide: vehicle attitude stabilization, solar collector pointing and figure control (currently passive for the photovoltaic satellite), and microwave (MW) antenna pointing and figure control, and stationkeeping in geosynchronous orbit.

Functional Requirements and Block Diagram

Functional requirements for various operating modes are listed in Table 3.1-16. The functional flow diagram in Figure 3.1-22 illustrate9 the major ACSS component subsystems and information flow between the components to satisfy the control and stationkeeping requirements. The ACSS is integrated with the IMCS which provides the interconnections for all the ACSS elements and the computational capacity for the control algorithms. The basic information for the implementation of the control laws is provided by the sensors. The control forces and torques are furnished by the ion bombardment thrusters of the reaction control system (RCS). The MW antenna pointing is achieved with the rotary joint and antenna gimbal torques.

Satellite Attitude Control Requirements. The attitude control system shall maintain vehicle stabilization and orientation accuracy in all three axes. The detailed performance requirements are given in Table 3.1-17. The coordinate systems used in the photovoltaic concept is shown in Figure 3.1-23. Attitude control RCS requirements as listed in Table 3.1-18.

Microwave Antenna Pointing Requirements. The MW beam steering is accomplished by a combination of mechanical antenna pointing and electronic beam steering. The mechanical gimbal pointing accuracy requirements must be > 1 arc-min. The antenna must be stabilized to < 1 arc-min/sec. The antenna figure control shall be capable of pointing each of the 34.9x30.6-m elements to an accuracy better than < 6 arc-min. The electronic steering of the MW beam to provide the vernier pointing accuracy is accomplished in the MPTS.

Stationkeeping. The purpose of the stationkeeping system is to maintain a geostationary equatorial orbit and spacing with respect to the other satellites

3-25

c Table 3.1-15. Design and Performance Characteristics

FACTOR

CONSTRUCTION SITE

CONSTRUCTION TECHNIQUE

MASS (KG)

MATERIAL

MAX ALLOW OP TEMP ("C)

OPERATING STRESS LEVEL (Wa) *

FACTOR OF SAFETY

MIN NATURAL FREQ (CYCLES/HOUR)

ORIENTATION

TOLERANCES (OUT OF PLANE) ABOUT Y AXIS ABOUT X AXIS

ANTENNA STRUCTURE

GE0

BEAM MACHINE

o.12x106

COMPOSITES

110/320

u "U cc Cn

1.5

2.0

NADIR

ko.8"

ROTARY JOINT STRUCTURE

GE0

BEAM MACHINE

0.6~10~

COMPOSITES

108

TBD

1.5

TBD

N/A

TBD

‘fU CC

= CRITICAL CRIPPLING STRESS; acn = CRITICAL BUCKLING STRESS

SOLAR ARRAY STRUCTURE

GE0

BEAM MACHINE

o.7x106

COMPOSITES

110

u =(5 CC cl-l

2.0

lO.O-LEO I,O-GE0

Y-POP Z-EQUATOR

0.3" 1.0"

J

Table 3.1-16. Attitude Control and Stationkeeping - Operating Modes

MODE

CONSTRUCTION

TRANSITION FROfl CONSTRUCTION OPNS

OPERATIONS

ECLIPSE

FAILURE MAINTENANCE

CHECKOUT

I I

i

f 1

FUNCTIONS

VEHICLE STABILIZATION DOCK1 NC STATIONKEEPING

REORIENT FROM CONSTRUCTION TO OPERATIONAL ATTITUDE

ATTITUDE CONTROLLED REFERENCE ORIENTATION

ANTENNA POINTING FIGURE CONTROL STATIONKEEPING

ATTITUDE CONTROLLED TO REFERENCE ORIENTATION

ANTENNA POINTING (STATIONKEEPING NOT REQUIRED)

FAIL-OPERATIONAL REDUNDANCY ON ALL ATTITUDE CONTROL FUNCTIONS

MAINTENANCE INTERVAL. > I YEAR -

LEAK CHECKS SOLAR POINTING AND FIGURE CONTROL STATIONKEEPING DYNAMIC RESPONSE

I BODY ATTITUDE

I

I SENSORS

4 * BODY

DYNAMICS l

SOLAR COLLECTOR 4 A POINTING

DISTURBANCE _, ENVIRONMENT

GIMBAL

I

+ 1 MW ANTENNA b I FIGURE SENSORS

Figure 3.1-22. Functional Flow Diagram

3-27

Table 3.1-17. Attitude Control Requirements

PARAMETER

PHOTO- VOLTAIC CR = 2

ASSEMBLY ORBIT GE0 .~ssy CONTROL (GRAVITY-GRADIENT z-Pop, y-L\

(STABLE) CONTROL ACCURACY (DEG) io.5

OPERATIONAL ATTITUDE CONTROL REFERENCE ATTITUDE Y-POP,X-IOF CONTROL ACCURACY (DEG) -co.1

CONTROL sys BANDWIDTH (CYCLESIHR) 0.5 SATELLITE FIRST BENDING MODE FREQ

(CYCLESIHR) 21.0

Y-POP x-top

Figure 3.1-23. Satellite Coordinate System

3-28

-- __..._..... _-..------ .._..-..,.,.. ---~

Table 3.1-18. Attitude Control RCS Requirements

in the presence of disturbing perturbations. These perturbation forces include the effects of earth gravitational anomalies, lunar and solar gravitational perturbations and the solar pressure force acting on the spacecraft. The thrusters that provide the forces and torques for attitude control also provide the necessary thrust for stationkeeping maneuvers.

The equatorial orbit is selected to minimize the impact of orbit inclination on rectenna size (and cost) requirements. This necessitates latitude (north-south) control.

The satellite longitude station must be selected within several degrees of its rectenna longitude in order to prevent an increase in rectenna size (and cost). The solar pressure induced perturbations are cyclical with an annual frequency and can be as large as k3.1' if uncorrected. In order to minimize the SPS space requirement in GE0 and to prevent interference with other satellites which do not experience as large a solar pressure perturbation as the SPS, it is assumed that this perturbation must be corrected. Because of the large magnitude of this correction, means of alleviating it should be investigated further in future studies.

To minimize interference with the large number of other satellites expected to be using this orbit by the 2000 time frame a stationkeeping accuracy of +O.l degree in longitude and latitude is adopted. The stationkeeping RCS requirements are summarized in Table 3.1-18. No stationkeeping thruster fir- ings should be performed during eclipse periods in order to minimize the thruster power requirements. Cyclic perturbations with a period less than or equal to one day need not be corrected.

3-29

Table 3.1-19. Stationkeeping RCS Requirements

Reaction Control System. The reaction control system (RCS) provides the necessary forces and torques for attitude control and stationkeeping. For the photovoltaic concept the RCS consists of four ion bombardment thruster modules with 16 thrusters at each corner of the vehicle. The argon propellant is stored cryogenically. A refrigeration system is ncess'ary to maintain the cryogenic temperatures. The thruster characteristics are given in Table 3.1.20.

Table 3.1-20. Electric Thruster Requirements 1

CHARACTERISTICS VALUE

THRUST 12 N SPECIFIC IMPULSE 13,OGO SEC PROPELLANT ARGON APERTURE 100 CM OPERATI tlG POWER 1275 K’d

Major Assemblies

Figure 3.1-24 illustrates the major assemblies comprising the ACSS. The description of each assembly, as applicable to the photovoltaic option, is given in the preceding section.


The point design ACSS is described in Volume II.

Subsystem Interfaces

The primary interfaces are the IMCS, the power distribution and control subsystem, and the structure. The IMCS, which functions as an integral part of the ACSS, also provides the interface for the ground support system to the ACSS. Figures 2.1-4 through 2.1-6 show the primary interfaces for the atti- tube reference system, the MW antenna pointing system, and the tank and engine system, respectively.

3-30

I

I 1

I ATTITUDE CONTROL 6 STATION KEEPING SUBSYSTEM

I

I I

I I 1 I

ATTITUDE M. W. ANT.

DETERMINATION FIG. REACTION PROPELLANT CONTROL CONTROL

: I SUPPLY

*FIG. SENSORS *THRUSTERS . TANKS . FIG. ACTUATORS . DRIVE l LINES

(31 ELEtTRONlCS . HEATERS . PROPELLANTS

C6LilMATORS . ELECTRONICS

I

. TORQUERS *ANGLE SENSORS

Figure 3.1-24.

3.1.6 THERMAL CONTROL

1 *VALVES

(1) r- -1 r-- i SECONDARY’ CRUCTUREA

--- -a- . INSTALLATION l Al-T.

DETERMINATION l FIG. CONTROL

(1) SEE STRUCT. SUBSY. (2)- SEE IMCS (3) ATT. REFERENCE SYS.

Assembly Tree - Attitude Control and Stationkeeping Subsystem

The thermal control subsystem continuously maintains temperature levels within allowable extremes and provides equipment for heat dissipation, acquisition and temperature regulation where required. Both active and passive systems may be employed and components utilized include selective coatings, insulations, heaters, radiator networks, and specialized energy transport devices such as heat pipes. Thermal control impacts almost all satellite operations supporting power conversion, power distribution, the microwave generator and power transmission systems, the rotary joint, information management, primary and secondary structural design, and the ground receiving station.


The thermal control subsystem must satisfy functional requirements during all satellite operating phases indicated in Table 3.1-21. A simplified functional flow diagram of the thermal subsystem, indicating its relationship to other operating subsystems is illustrated in Figure 3.1-25. The klystron radiator heat pipe assembly is shown in Figure 3.1-26.

3-31

Ta Xode I Assemblv I Cunccions ^" --I

I . . :onsfruccion 1 Subsystem

:nter-Orbit Transfer Subsystem

)peracions Subsystam

I 1

'la-ntain ailcwable tempeiatl~re irveis and gradients tc restrict structural deformations/stresses and protect assemblies. I Gineain allowable temperature levels and gradients; power required operation phases. 1

1 Support steady-state operation for all assemblies.

Le 3.141. Thermal Control Subsystem - Operating Modes

1 Klystron Radiators/ Restrxt thermal stresses which could : impan microvave transmission. .Xin- imize heat transfer to rotary joint, antenna structure, and electronics' modules.

Guarantee integrity of all systems A

ClipS during extended cooling and return cc : Subsystem steady state; assure ccntinucus

cperacicn of resumption after shutdown, as required. I

Drain fluid as necessa? and provide P/C* Radiators localized heating if required: smooch I

res.cart to steady-state operation. I

Recover from fluid (hear pipe) freeze-' Klystron Radiaccrs uD

I e.;.. hear oises. mm&. Provide Reaundanc capability where Possible: '

' rapid access to down components.

Leak isolation through application of :

fication by sensors, or possibly by

leckout Subsystem Ground test vhere possible: leak checkj verify control response. J

?/C - Power Conversion Subsystem I

TEMP LEVELS

d TEMP LEVELS ANTENNA

t

OPERATION

NORMAL OPERATION ECLIPSE

PERTURgATIONT

SHOV8RUSH TEMP LEMLS

ACTIVE WASTE HEAT REJECTION SYSTEM

STRucTUP.AL THERMAL DIST L TEMP LEVELS

CONDUCTOR TEMP

CONVERSION

NORMAL OPERATIONS

PERTURBATIONS

Figure 3.1-25. Thermal Control Functional Flow Diagram

3-32

KLSTRON COllECTOk

Figure 3.1-26. Klystron Radiator Configuration

Major Assemblies

Figure 3.1-27 illustrates the major assemblies and components comprising the thermal control subsystem.

Designand Performance Characteristics

Design and performance data for the klystron radiators are presented in Table 3.1-22.

3.1.7 INFORMATION MANAGEMENT AND CONTROL

The information management and control subsystem (IMCS) provides the interconnecting elements between and within all the various satellites and ground-based operational subsystems. The IMCS also provides operational control of both the satellite and ground systems as well as providing all subsystem processing support for all but very special functions.

The satellite IMCS consists of the on-board processing equipment [central processing units (CPU) and memories], the inter- and intra-subsystem data network (data buses), the man-machine interfaces (display/control), and inter- system communication links, including RF, but excepting those specifically provided for the control and transfer of primary power, and all elements provided to accommodate activities related to system security, safety, or any other operation necessary to the continuing operation of the SPS.

Because of the early stage of program analysis, only those requirements imposed upon the MCS by a limited number of satellite operations have been identified. The identified requirements generally are limited to those associated with the immediate operations of an active satellite. Auxiliary functions such as ground/space communications, display/control, safety, security, etc., will be added when data become available.

3-33

THERMAL SUBSYSTEM

I

POWER CONVERSION (PHoToVOLTAIC)

‘COATINGS l SENSORS

I ------e--m--- 7

1 1 r ----A -w-m ‘I

ANTENNA ROTARY I I SYSTEMS JOINT I RECTENNA

IN-.----.- - - - 1

l COATINGS . COATINGS l COATINGS . INSULATION l SENSORS . KLYSTRON HEAT PIPES .FINNED CONTAINERS

FOR ELECTRONICS . SENSORS l HEATERS

(1) r

------- 1 I SECONDARY 1

L STRUCTURE -1 ----m-e

. COATINGS (I) MAJOR INTERFACE

Figure 3.1-27. Assembly Tree - Thermal Subsystem

!

Table 3.1-22. Klystron Cavity Radiators (Slaximum Intensity Region)

Total heat load (kW) 3.264 Driver cavities (kW) 0.206 Output cavity (kW) 2.308 Electromagnet (kW) 0.750

Radiator temperature ("C) Radiator area, mine (m2> Fin material Fin .efficiency (X) Coating (external) Coating (internal)

198 2.36 Aluminum 60 Anodize (soft) Anodize (hard)

Heat Pipes - Four high-performance, arterial wick copper/water heat pipes of 28-in. length each l/2 in. O.D. Twenty-eight axial groove copper/water heat pipes, 25 inches long, each 3/B in, O.D. (Container is actually copper liner encased in aluminun tube). Total heat pipe assembly weight = 6.18 kg.


The functional requirements for various operating modes are given in Table 3.1-23. The relationship of the IMCS to the other major subsystems is depicted in Figure 2.1-l. Figure 2.1-10 illustrates a representative processor hierarchy as applied to a solar photovoltaic power-generating satellite concept. The IMCS hierarchy applicable to the microwave antenna subsystem, attitude control and stationkeeping subsystem, and power distribution subsystems is presented in Figures 3.1-28 through 3.1-30, respectively. These hierarchies are established to the level at which the IMCS and the using subsystem interfaces are apparent (e.g., physical/electrical interface).

Table 3.1-24 summarizes the estimated number of data interfaces (not measurements) that must be accommodated by the IMCS. Note specifically that the microwave antenna subsystem is by far the major contributor to the determination of the complexity of the IMCS electrical interface. Table 3.1-25 provides a very preliminary estimate of the control interface that must be accommodated by the IMCS although the estimates for the other subsystems are not supported by an in-depth analysis. Again, the microwave antenna system predominates.

Major Assemblies

Figure 3.1-31 identifies the major assemblies that form the IMCS. six major assemblies have been identified at this time:

control units (BCU), (3) data bus, (4) (1) processors, (2) bus

remote acquisition and control units (RAC), (5) submultiplexers (SM), and (6) microprocessors (up).

3-35

Table 3.1-23. IMCS - Operating Modes

Xode

Conrtluction

Inter-Orbit Transportation

operations

Eclipse

Transition

FaLlurelHaintenance

Assembiv

Subsyse‘em

Subsysteu

Subsystem

Subsystem

Sobsys tern

runct1on \

Temperature monitor Attitude monitor and control Safety monitor

Power conversion and distrib.xtion Nonitor and control Navigation Attitude monitor and control Subsystem monitor Configuration control MW pointing, gimbal pointing

control

Steady-state monitor and control

Eclipse monitor Shutdown/startup monitor and

control Subsystem standby monitor and

control

Orjentacion monitor Subsystem monitor and control

Failure detection/isolation Redundancy management Auto shutdown/restart Override control Kaintenance logging

Processors. The satellite Master Control Computer (Figure 3.1-28) will operate with a 16-32 bit word format and have a 64K-128K word active memory plus a TBD billion word bulk storage facility. Second- and third-level processors (supervisory or local) will be 16-bit word assemblies and be limited to 16K-32K memories. In special cases, memory capacity may be increased to as much as 128K words. Assemblies or subassemblies identified as microprocessors (normally those units incorporated directly within the associated electronics) will incorporate an 8-bit-wcrk format and use active BK-64K word memories.

Bus Control Unit. The bus control unit (BCU) provides the control necessary for data/command transfer over the subsystem data bus network. The BCU accepts instructions and data (or commands) from its associated processor and translates these data from a processor-compatible format to one compatible with the data network. It also accepts bus-compatible data and converts these data to processor formats. In addition, the BCU monitors the data traffic--performing bit and word checks as well as health/status checks.

In addition to data bus control, the BCU will provide a computer-to-computer link where appropriate.

Data Bus. The data bus network accommodates multiplexed, digital data transmitted between the BCU and all other remotely located data acquisition and contra1 devices associated with a specific processor/BCU combination. The bus link may utilize conventional wire techniques for short runs in low EM1 areas or fiber-optic technology for long paths or through high EM1 areas. Basic bit rate within the bus assembly is assumed to be 1.0 Nbps. Included in the data bus assembly are the data bus coupling devices used to connect the various

3-36

MASTER

CONIXOL

COMFulfR

Ku

,

c

10 OTHR slJrERvIsow CoMNTfIS

MlaowLvl *NltNN*

suffRvIsonY COM;WCR

KU

huaow*vT

ANTINNA c COMnmR I

Ku

I MIaOWAVf *NltNN*

COmFlJltR 3 x5J

!

l&P _ P I 75

Ku 1 Ku

IAC - ItMOlf ACQlJISIlION MD CONlROL UNll

SM - suI-MJllInfx~n

pr - MlaO-mOCESSOI

IOJ -US. CONlKOL UNIT

Figure 3.1-28. IMCS - MW Antenna

I I . . RU UC

31’ . . . I 1

Figure 3.1-29. IMCS - Attitude Control

WTEl CmJnoL

DISPIA~ .-

COMnllR * AND

KU CONlIOL

cwm SfclIoN I LlGwmG-

I HXIIP*YNT

r- ------ J I I ToorHm

I I

I

I stlmvlsoIY

fOHR DlSl’RlblJlIoN ’

f

I CoMNlES POWS DISRIUJllON

sUMMsoIY SUNRVISOW COYnJlR cohvu1n

4 KU

I I

‘Ku

I I

I I . KNEI c4SnlmJTloN

CINtR

CDHPulR KU \

I fOWER DISlRIllJlION

RIGHT WING

COMRLIU ‘2 KU

RICH1 WlNG

i I

!

j

i

i i i : : I :

I i I

I E

r : I i ! I i I i

I i i

Figure 3.1-30. IMCS - Power Distribution

Table 3.1-24. Preliminary Data Interface Summary - Photovoltaic (CR-2) Configuration

ANALOG DIGITAL EVENT TOTAL

HI CROWAV E ANTENNA 6x10~ 1X106 2. 1X106 >gx 1 o6

OTHER SUBSYSTEMS STRUCTURE

92 I2 35 >lOO

ATT. CONTROL 8 STATIONKEEPING 1000 ..3000 POWER DISTRIBUTION 1000 100 2000 -3000 INFORYATION HANAGEHENT 719,000 -I9,000 THER%L 16,000 16,000 LIFE SUPPORT TBD Ti3D TBD TBD SAFETY AND SECURITY TBD TBD TBD TBD

Table 3.1-25. Preliminary Control Interface Summary - Photovoltaic (CR-2) Configuration

HI CROWAVE ANTENNA

OTHER SL’SSYSTEHS STRUCTURE ATTITUDE CONTROL & STATIONKEEPING POWER DISTRIBUTION INFORYAT ION MANAGEMENT THERML LIFE SUPPORT SAFETY AND SECURITY

PROPORTIONAL

<13.6x104

TBD TBD

EVENT

3oxto4

-35 >300 >300

>3000

Cl00 c500 >300

>3000

TBD TBD TBD TBD

TOTAL

<44x104

I BUS RLWJIE

PRuCESSOR CONTROL ACQUISITION L CONTROL

11.x2

l H*IIDWARE - CPU - MEMORY

* SoFlwARE

-DATA WS

- I

*HARDWARE COUPLER - CPU

l IRIDGE - MfMOKY

COUPLER l SO+TW~L l fI8ER o*TIcs

0 r--- “1 t , COMb!OL :

L-,--l

l fUNCTIONS l ffOCEDURES

(1) r-m -w--q I sECONJAxY ’ I STRUCTURE 1 L s----- J

* IbiSTUUllON

(1) SEE STRLKTURAI. SUBSMTEM 0 SLE APP:lCAKE SUSSnTEJ.6

Figure 3.1-31. Assembly Tree - Information Management and Control Subsystem

3740

remote units (one required per remote) as well as the bridge coupler required to transfer data across the microwave antenna rotary joints; the latter element is presently TBD.

Remote Acquisition and Control. The remote acquisition and control (UC) assembly is the basic interface between the IMCS and the various operating subsystems. The RAC provides for data format conversion from the preconditioned analog, digital or event voltage/impedance levels, and converts these data in- to 8-big digital, serial, equivalents. The RAC also accepts digital data words and outputs commands in a format compatible with the receiving subsystems.

Basic conversion (input/output) is assumed to be fl% (e.g., 7-bit and sign). Voltage ranges and impedances are TBD.

Submultiplexers. The submultiplexer (SM) provides a means of expanding the capability of the RAC. The SM thus contains all of the capabilities of an RAC, but can only communicate with a single UC rather than a given data bus. The number of SPI's that can communicate with an RAC is presently TBD.

Microprocessor. The microprocessor (up) elements provide local, front- end processing of data obtained from the various using systems. These processors will handle the bulk of the system's monitoring and control task, sending raw data up through the computer hierarchy only when the task-levels exceed preestablished limits, or when detected out-of-tolerance conditions exceed local control boundaries. These devices are solid state and could normally be integrated within the user electronics. When necessary, the pp can be located within the RAC's or SM's to provide local performance monitoring and control.

Subsystem Definition and Interface

The subsystem interfaces for the three major subsystems are indicated in Figure 2.1-2 through 2.1-11. Table 3.1-26 summarizes the number of IMCS elements required for a typical photovoltaic configuration. Table 3.1-27 summarizes the physical (weight, power, volume) requirements for this system.

3-41

Table 3.1-26. Hardware Summary

fUUCTlON -

‘ROTE MASTER DISPLAY SUPER- IUS muIs.

HARDWARE CONTROL AN0 VlSORY REnOTE HICRO- CONTROL AND ELEMENT COMPUTER COnTROL SUJ-MUX.

SATELLITE 2 I m w - 2 - - CONTROL

THERMAL - - CONTROL 2 5 m 7 a5 1.352

STRUCT. - I . - - - ALIGN. 3 3

ATTITUDE CONTROL 1 10 I1 2a 148

POWill OISTRI6. . 2 5 I 7 as -

HI CROYAVE ANTENNA c 1 14 777 792 787 29,500 CONTROL

TOTAL 2. 1 6 37 777 I

622 I

985 31,000

Table 3.1-27. Weight/Power/Volume Summary - IMCS

ION.ROlATlNG

I UNIT UASS

hARDWARE ELEMENT 1 OUANTITY 1 IKpl

YL;STER CONTROL COMPUTER I 2 I 500 DISPLAY 6 CONTROL SET 1 200 SUPERVISORY COHPUTER 9 14 REMOTE COVPUTER 23 14 YlCROOROCESSOR 5 IUS CONTROL UNIT 1 5 RCM~~EACO~JISIT~ON~ CONTROL 1 ,l9l9 SUE ~ULTIPLEXOR

SUETJlAl

: 1

1 1.232

150 990

L SO0

IOTATI!dG

UAf7ER CONTROL COUPUTER I DlSPLAY 6 CONTROL SET SUPERVISORY CO*.IPUTER REMOTE COVPUTER

I 1 1 14

MICRO PROCESSOR 111 BUS COSTAOL UYIT 792 REMOTE ACOUISITION 6 CONTROL 767

TOTAL MASS II91

l.COO 206

70 322

UNIT i TOTAL i UNIT VOLUME

lm3t

0.4 0.9 0.) 0.72 0.07 0.35 0.01 _I ] 0.07 I 1.61 0.01 0.02 0.003 0.02 0.6 0.005 0.02 3 96 0.005 0.99 001 15.0 0 003 45

26.42 , 7.44

2 09 0 07 0 07 0.02 0.02 0.02

- --~ 0.4 0%

SUE VULTIPLEXOR

CABLE

YON~ROlATING-WIRE 122GAI 1.200 KM 12 O/KM 14.000 FIBER OPTICS ?9 KY 0.141KM 12

ROTAllnC-WIRE 23.000 KM 279,000 Fl6ER OPTICS 350 KM 50

to,., / 191000

3-42

3.2 GROUND RECEIVING STATION

The following subsections of this document describe the requirements, major assemblies, characteristics, an'd definitions for the subsystems comprising the ground receiving station(GRS) element of the SPS. An artists illustration of the ground receiving station complex is shown in Figure 3.2-l. The major effort to date has been limited to the establishment of the receiving/ rectifying portion (rectenna) and the power distribution network. A limited evaluation and characterization of the data management and control subsystem and the data conversion system has been made. A final area lightly touched upon during the course of the study was a preliminary identification of the need for a separate beam monitoring system to backup the retrodirective beam concept. No data for the latter system has been derived. The assembly tree for the overall GRS is shown in Figure 3.2-2.

STORAGE 6 MAINT.

USER POWER TRANS.

ERTZ)

A --

MONITOR 6 COMROL FACILITY

KMX 13

Figure 3.2-l. Operational Ground Receiving Facility (Rectenna) - Typical

A separate study activity, under task 2 of the primary SPS study, was made to evaluate the system control requirements. The results of this latter study is documented in Section 8.0, Volume V of the final report.

3.2.1 RECTEXNA

The rectenna subsystem consists of microwave receiving elements (dipoles), rectifiers, regulators and isolating motorswitches (Figure 3.2-3). The dipoles are fabricated using a multilayer (sandwich) construction of copper and dielectric insulators formed into panels. A rectification element consisting of a GaAs diode and filters is added to convert the received microwave energy into dc. Conversion efficiency is estimated to be 89%.

3-43

GROUND RECEIVING STATION

I I I I I 1

POWER RECTENNA DISTRIBUTION POWER THERMAL

AND CONTROL CONVERSION CONTROL

DATA SAFETY STRUCTURES COMMUNICATION MANAGEMENT AND

AND CONTROL SECURITY I

Figure 3.2-2. Ground Receiving Station Subassembly Relationships

r I I I 1 ANTENNA RECTIFIERS REGULATORS SWITCHING

- DIPOLE CONFIG. - DIELECTRIC * COATINGS * FABRICATION * MODULARITY - SENSORS - INTERCONNECTS

* DIODES * TYPES * FILTERS * LOCATION - MATERIALS * MATERIALS * SENSORS ’ SENSORS . IHTERCONNECTS ’ INTERCONNECTS

* MOTOR SWITCHES . SENSORS - INTERCOhNECTS

Figure 3.2-3. Assembly Tree - Rectenna


The functional requirements for the rectenna subsystem are listed in Table 3.2-l. A simplified schematic block diagram is presented in Figure 3.2-4.

Major Assemblies

The major assemblies and components that are required for the rectenna subsystem are shown in Figure 3.2-5.

Antenna. The antenna is a multilayer copper/dielectric sandwich panel as shown in Figure 3.2-5. The total antenna system consists of 580,500 panels each 9.33X14.69 m. These panels areinturn made up of twenty 0.74x9.33 m sub- panels mounted on a supporting structure (see structure subsystem). Total surface area (in GRS) is 79.56 km'.

3-44

Table 3.2-l. Rectenna Functional Requirements

'ROGRAMMATIC ENERGY SOURCE - HICRO!4AVE AT 2.45 GHz CAPACITY - 5 GW (NOIIIIIAL) DELIVERED TO ~tiz DISTRIBUTION NETWORK LIFETIdE - 3G YEAAS WITH MlN;tll;M PLAHIIEC t’4 , I’4TEN;rNCE (S:-IO!JLD CE CAPASLE CF EXTEE!DED LIFE BFYOND

30 YEARS WITH REPLACEMENT) IOC DATE - 2000 OPERATIONS - ANYWHERE WITHIN OR IMHEDIATE~Y ADJACENT TO CONTINENTAL U.S.A. RESOURCES - tllNlHUH USE OF CRITICAL RESC,?.,ES COMt4ERClALlZATlON - COHPATl8LE WITH UNIT:: STATES UTILITY NETWORKS DEVELOPHENT - EVOLUTIOZARY, WITH PROVISlC’.S FOR INCORPORATING LATER TECHNOLOGY

TECHNOLOGY

OUTPUT POWER-POWER LEVEL IS DEFINED AS C:‘;STANT POWER LEVEL (5 GU, MAX), EXCEPT DURING SOLAR ECLIPSE

ENERGY STORAGE-NONE FAILURE CRITERIA-NO SINGLE-POINT FAILURE “:Y CAUSE TOTAL LOSS OF SPS FUNCTION ENERGY PAYBACK-LESS THAN THREE YEARS COST-COMPETITIVE WITH HYDROCARBON OR HY::::LECTRIC POWER GENERATION CONCEPTS WITHIN LIFETIME

OF SPS PROJECT

MODE CONSTRUCTION OPERATIONS

ECLIPSE

FAI LURE,“fiH~TENANCE CHECKOUT

9PERATION

ASSEMBL'! FUNCTION

SUBSYSTE” NONE SUBSYSTE’ STEADY-STATE OPERATION

SUBSYSTE” OPEN ISOLATION SWITCHES CLOSE ISOLATION SWITCHES

SUBSYSTE’ VOLTAGE CHECKS: SWITCH STATUS SUBSYSTE” FAIL-SAFE CHECKS; CONTROL RESPONSE

ANTENNA ’ INPUT FILTER ’ DIODE ’ OUTPUT FILTER

Figure 3.2-4. Simplified Schematic - Rectenna

Rectifier. The rectifier assembly consists of a GaAs diode and input/ output filters. An illustration of a possible diode configuration is shown in Figure 3.2-5. The equivalent schematic of the rectifie;/filter circuit is shown in Figure 3.2-4. The outputs of the rectifier circuit are series connected to output 40+ kV.

Regulators. The regulation assembly accepts the voltage from the series . . . . . ._ connected rectenna diodes and adjusts the voltage output to the power distribu-

tion feeders to a value consistant with positive current flow.

3-45

Cu Clad Mylar

Dielectric

Hard Su’face>ARRAY ‘CROSS SECTION

( im) tiAS GaAs

SUBSTRATE’ EPIIAYER .\

\J------- Au (2 I LEAKAGE 5 A AT 21C”C

Im’Y)mA-38lmA

, PtCr OR NICHROME MXtDA)

V,.O.lV AREA - 10-3 CM2

COULD ALSO BE SNAP-ON HOWEVER, SNAP-ON HAS SERIOUS DEGRATldN PROBLEhIS - ESPECIALLY

‘(111) QAs IS ALSO A CANDIDATE WITH MOISTURE.

FOR SCHOl-fKY BARRIER DIODES

DIODE CONCEPT Figure 3.2-5. Rectenna Systems Major Assembly/Component

Switching. The motor switches provide for no load isolation of independent voltage string.

each


The design and performance characteristics of the.rectenna subsystem are presented in Table 3.2-2.


The subsystem interfaces are shown in Figure 3.2-3. Details of the interface are TBD.

3.2.2 POWER DISTRIBUTION AND CONTROL

The power distribution and control subsystem receives power from the rectenna subsystem and provides the switching required to deliver the power to the power conversion stations, and then delivers the power station outputs

3-46

Table 3.2-2. Rectenna Preliminary Specifications

I TER CHARACTERISTIC

INTERCEPTED ENERGY (GU) FREQUENCY (GHz) z: RECiENNA EFF I C I ENCY (%) as CLUSTER OUTPUT TBD VOLTAGE STRING OUTPUT (KV) 40+ RECTENNA OUTPUT ENERGY (GW) 4.93 NUMBER OF DIODES 330X106 RECTENNA SUBPANEL SIZE (tl) 0.735x9.33 PANEL DlttENSIONS (ti) 14.69x9.33 NUMBER OF PANELS IN RECTENNA 580,500 PANEL AREA (Ml’) 79.56 RECTENNA CDNFIGURATION ELL I PSE RECTENNA DIMENSIONS (KM) 10x13 RECTENNA GROUND AREA (KJI’) 102.5

to interconnected utility interfaces. The feeders, and power cabling as well as internal transmission towers and cabling are included. Power transmission, (high tension cabling), ,from the designated interface at the perimeter of the ground receiving station are the responsibility of the power utility. The grounding, electromagnetic interference control, and all shielding requirements are also included. The life expectancy of the power distribution system is 30 years. The responsibility for auxiliary power systems used to maintain critical subsystems is TBD.

Functional Requirements and Diagrams

Functional requirements for various operating modes are listed in Table 3.2-3. A specified schematic block diagram for the ground receiving station is presented in Figure 3.2-6.

Table 3.2-3. Power Distribution and Control - Operating Modes

I MODE 1 ASSEHRLY 1 FUNCTION I-

CONSTRUCTION

OPERATION

ECL I PSE

N/A

SUBSYSTEM

SUBSYSTEM

N/A

STEADY-STATE OPERATION

STARTUP/SHUTDOWN, BACKUP POWER TO CRtTtCAL SUBSYSTEtIS

FAILURE/ &AlHTENANCE

CHECKOUT

SUBSYSTEM

SUBSYSTEM

REDUNDANT OPERATION. AUTO SHUT- DOWN

CONTINUITY, INSULATION RESIST- ANCE SWlTCHtNG RESPONSE

Major Assemblies

Figure 3.2-7 illustrates the major assemblies comprising the power distribution and control subsystem.

Power Distribution. The power distribution assembly consists of the main feeders, secondary feeders, 40 kV dc and 500 kV ac buses, tie bars and power interface cabling for the various operating subsystems. The main feeders are sized to handle gradually increasing current loads starting at the center of

3-47

L

96 DOUBLE BOXES EA SECTOR 1M . CUBES

\ 40,000 VDC BUSSES

’ ‘W 14.8 MW

ISOLATION

3 EMSECTOR

_------------------ FEEDER POWER (INPUT TO 40,ow VDC SUSSES) -

RANGE 12.1 - 16 MW AT (300 - 400A) AS MANY FEEDERS

AS REQUIRED TO MASS IN QUADRANT (FEEDERS ONLY) SO,ooO KG (200,~ KG FOR ENTIRE FARM) APPROACH 16 Mw

Figure 3.2-6. Rectenna Schematic Block Diagram - Preliminary

\ ’ 1088 ROWS

POWER DISTRIBUTION AND CONTROL

I I I- I

POWER SWITCHING I DISTRIB, I

l- - ----

. MAIN FEEDERS . GEAR . 40-KV DC TOWERS

. SECONDARY . SOLID STATE . 500-KV AC TOWERS

FEEDERS . SENSORS . INSTALLATION . BUSES . SfNSORS . IN\tAlCATION

Figure 3.2-7. Assembly Tree - Power Distribution and Control

3-48

the rectenna array and continuing to the perimeter. The feeders are grouped in each quadrant of the array to permit systematic maintenance and to avoid catastropic system failures. The main feeders utilize TBD cm round aluminum cables, uninsulated, mounted on insulated standoffs or in insulated raceways. Other feeders, tie lines, buses, etc., are sized to handle maximum estimated loads, at specified voltages. All cables are passively cooled by radiation to local environment.

Switching. Switchgears are used for:

l Isolation/selection of various power feeders as a result of changes in power demand or as the result of systematic element failures.

8 Isolation/selection of power conversion stations as load demand varies or due to systematic failure.

l Isolation of loads as satellite power capability varies due to predicted (eclipse, maintenance, etc.) power reductions or due to unpredicted (systematic failures) power reductions.

The switchgears may be solid-state or electromechanical. The voltages and currents being handled by these switches will be monitored by the IMCS to determine their status and to establish a need for the automatic opening of these switches (circuit breaker function). Switch closure will be based upon fault status and power demand. During shutdown operations the system will be monitored and when certain conditions are reached a command signal will automatically open or close selected switches as required.

Control. The power distribution control concept is based upon a continuous monitor function performed by the station resident IMCS. The IMCS also formats concise system display summaries to permit efficient transfer of information to the system operators. Where control discussions must be made at a rate beyond that possible through human intervention, preprogrammed control sequences will be initiated to establish desired system configuration. Primary system control, except for emergency situations, is vested in human operators.

Included in the general category of control are the functions associated with the man-machine interface, i.e., display and control.

Secondary Structures. Secondary structures consist of mounting brackets, clamps, raceways, as well as all other secondary installation devices as needed. It is assumed that a delta of TBD percent of the subsystem mass is reason- able.

Transmission Towers. The 40 kV dc and 500 kV ac power buses are supported by suspension towers around the perimeter of the rectenna area but within the outer station perimeter fence. The 40 kV dc supports consists of four 18 meter high, tapered, steel poles. The 500 kV ac towers are standard 70 meter towers similar to those used for cross-country transfer of power from sources such as Hoover Dam or TVA.

3-49


The design and performance characteristics for the power distribution subsystem are listed in Table 3.2-4.


Halor Assembly I Keqllireme"ts Technology Ishug. re

( :ENERAL Hllss HTBF Life Efficiency Resupply and maintenance

POWER DISTRIBUTION (PD) Mass Material Ipsulation Efficiency Subsystem cabling Resupply and maintenance Life

3w r'rctt (:I:AZH Ih-lla I1y ‘rypc I’lwur rc11 I nv. VII I t ;&xc- El I II, Iwwy I.ll’C l~csupply and m;llntenance

Configuration dependent Subsystem dependent 30 ycors 88-98% (config. dependent) As needed

Mostly round conductor Conf igurot ion dependent Aluminum 6001-T6 TBD 88-98X (config. dependent) Locat ion and power dependent As rcquircd 30 years or greater

Al’lwux. 0. DOlJtlh kJ;/kW SatI Id .ut:lI~’ (:011i l J:llWl ~1111 Ihhpl’ll~hl~ (:I1111 lg. i,ll,l Illl.ilL 1,lll dvl”~ll’I’wL ‘J’l-‘VI . ‘J% III Y(*:,I’H

As rPl(l! I red

SECONDARY STRUCTURE Mass TBD% of PDS weight was considered

to be required for mounting and installat ion.

CONTROL Temperature sensors Current sensors Voltage sensor‘5 Switch gear cor.trol

Overcurrent Overvoltage Undercurrent Undervol tage

No. of sensors config. dependent No. of sensors config. dependent No. of sensors config. dependent conflguratfon dependent


Subsystem interfaces are shown in Figure 3.2-6 for the power distribution subsystem approach selected for ground receiving station. Power handling capacity is estimated to range up to 5.0 Gtl.

3.2.3 STRUCTURES

The GPS structure assemblies considered in this report are primarily those associated with the support of the rectenna panels, plus the secondary elements already discussed in Section 3.2.2. Included in this subsystem are concrete footing, steel primary and secondary support structure, bracing and the various

3-50

-

4.0 SUPPORT SYSTEMS

4.0 SUPPORT SYSTEMS

4.1 GE0 OPERATIONAL BASE

0-D)

4.2 MAINTENANCE AND REFURBISHMENT FACILITY

(TBD)

4.3 SPS TRANSPORTATION SYSTEM REQUIREMENTS

The SPS program will require a dedicated transportation system and, in all probability, a dedicated launch facility for vertical launch HLLV operations.

The major elements of the SPS transportation system consist of the following:

. Heavy-Lift Launch Vehicle (HLLV)-SPS cargo to LEO

l Personnel Launch Vehicle (PLV)-personnel to LEO (Growth STS)

l Electric Orbit Transfer Vehicle (EOTV)-SPS cargo to GE0

l Personnel Orbit Transfer Vehicle (POTV)-personnel, LEO to GE0

l Personnel module (PM)-personnel carrier, earth to LEO to GE0

l Intra-Orbit Transfer Vehicle (IOTV) -on-orbit cargo transfer

Two HLLV configurations are required- a two-stage vertical takeoff horizontal landing (VTO/HL) HLLV with a payload capability in the order of 225,000 kg for the operational program, and an interim Shuttle transportation system (STS) derived HLLV for precursor operations. The latter vehicle utilizes the same elements as the PLV except that the orbiter is replaced with a payload module and a recoverable engine module.

The PLV is used to transfer the SPS construction crew from earth to LEO. This vehicle is a growth Shuttle version in which the solid rocket booster (SRB) is replaced with a reusable liquid rocket booster (LRB). The PM is designed to fit within the existing orbiter cargo bay.

The EOTV is employed for cargo transfer from LEO to GEO, and utilizes the same power sources and construction techniques as the SPS. The configuration, payload capability, and trip time are established on the basis of overall SPS program compatibility.

4-l

The POTV is the propulsive element required to transfer the PM and its crew/passengers from LEO to GEO. The POTV is a single, chemical rocket stage and is sized to fit within the cargo bay and payload capability of the PLV.

The PM is capable of transporting a 60-man construction crew from earth to LEO to GE0 and return. The PM is also sized to fit within the PLV payload envelope.

The IOTV, defined in concept only, is a chemical rocket stage, manned or remotely operated, and is capable of on-orbit transfer of approximately 225,000 kg of cargo over a distance of 10 km.

4.3.1 TRANSPORTATION SYSTEM SCENARIO

Transportation system LEO operations are depicted in Figure 4.3-l. STS derivatives are employed for crew transfer from earth to LEO. The STS-HLLV is employed early in the program for space base and precursor satellite construction and delivery of POTV propellants. This element of the operational transportation system is phased out satellite construction, or sooner. cargo and propellants to LEO, which the IOTV for subsequent transfer to

of the program with initiation of first The SPS HLLV delivers operational phase are transferred to the EOTV by means of GEO.

LEO STAGING EON TO GE0

Figure 4.3-l. SPS LEO Transportation Operations

4-2

Transportation system GE0 operations are depicted in Figure 4.3-2. arrival at GEO,

Upon the SPS construction cargo is transferred from the EOTV to the

SPS construction base by IOTV. The POTV with crew module docks to the construction base to effect crew transfer and POTV refueling for return flight to LEO. Crew consumables and resupply propellants are also transported to GE0 by the EOTV.

Figure 4.3-2. SPS GE0 Transportation Operations

Transportation system requirements are dominated by the vast quantity of materials to be transported to LEO and GEO. Tables 4.3-1, 4.3-2, and 4.3-3 summarize the mass delivery requirements, the baseline transportation elements.

and numbers of vehicle flights, for All mass figures include a 10% packaging

factor. Table 4.3-l summarizes transportation requirements for construction of the first satellite. Table 4.3-2 is a summary of requirements during the total satellite construction phase (i.e., the first 30 years). The average annual mass to LEO during this phase is in excess of 130 million kilograms with more than 750 HLLV launches per year. Table 4.3-3 presents a total program summary through retirement of the last satellite after 30 years of operation. Mass and flight requirements are separated between that required to construct the satellites and that required to operate and maintain the satellites.

4-3

Table 4.3-l. TFU Transportation Requirements

SATELLITE CONST. MAINl. 8 PACKAGING

CREW CONSUMABLES 8 PKG,

POTV PROPELLANTS 8 PKG,

EOTV CONST,, MAINT, 8 PKG,

EOTV PROPELLANTS 8 PKG. IOTV PROPELLANTS 8 PKG.

TOiAL

TFU FLEET

MASS x 106 KG

LEO GE0 .

37,12 37012

0,98 0.94

2.91 1.46

7020 - 4.79 - 0.13 0,06

-- 53813 39,58

GROWTH SHUTTLE VEHICLES-

PRECURSO: REQUIREMENTS: *LEO BASE *SPACE CONSTR, BASE l EOTV TEST VEHICLE

PLV HLLV

45

15

60

163.5

4.3

12,8

31.7

21.1 046

- 234-O

HICLELEF.LIGHTS i=

POTV

45 6.5

0,2

0,3

235 ,-

Ic LEO - 4

164

4

13

32

21 1

I GE0

164

4

6

VEHICLE REQUIREMENTS

2 15 14 6 4

PERSONNEL (PLV) CARGO CARRIER/ENGINE MODULE AND LAUNCH VEH

72 FLIGHTS 129 FLIGHTS 1 VEHICLE 2 ~. ~. VEHICLESm.-.-

Table 4.3-2. SPS Program Transportation Requirements, 30-Year Construction Phase

MASS x lo6 KG VEHICLE FLIGHTS PLV HLLV POTV EOT'J IOTV

AEn r,E@ LF" GE0

SATELLITE CONST. 8 MA1NT.n 3,099,3 3.099.3 3187 13.653 3051 599,5 13.653 13,653 CREW CONSUM.4BLES 74.9 71.7 - 33cl - 13.9 330 316 POTV PROPELLANTS 21606 108.3 - 954 - 20.9 954 477 EOTV CONST, 8 MAINTENANCE 38.4 31,2 - 169 - 6.0 169 137

EOTV PROPELLANT 492.3 2,o - 2.169 - 0.4 2.169 9

IOTV PROPELLANT 10,5 408 - 47 - ] o.gl-TT2/-+

TOTAL 3,932,O 3,317,3 3187 17,322 3051 642 31,935

VEHICLE FLIGHT LIFE 100 300 100 VEHICLE FLEEl' REQUIREMENTS - - 32 58 31

I 20 200 32 160

4-4

Table 4.3-3. Totdl Transportation Requirements, 60-Year Prog.ram

ll I i, MASS x IO6 KG I VEHICLE FLIGHTS r

POTV EOTV IOTV LEO 1 GE0

SATELLITE CONSTRUCTION 1220 425.1 9682 9682 OPERATIONS & YAINTENANCE 3660 348.7 7943 7943

CREW CONSUMABLES CONSTRUCTION 139 c 5.6 139 126 OPERATIONS & YAINTENANCE 86.8 86.0 - 382 - 16.6 382 379

POTV PROPELLANTS / I

CONS'lWJCTION 82.7 41.4 - 364 - 8.0 364 182 OPERATIONS & MAINTENANCE 267.6 133.8 - 1180 - 25.9 1180 589

EOTV CONSTRUCTION CONSTRUCTION 28.2 24.2 - 124. - 4.7 124 107 OPERATIONS & MAINTENANCE 22.2 19.0 - 98; - 3.7 98 84

EOTV PROPELLANTS CONSTRUCTION 340.3 2.0 - 1499 - 0.4 1499 OPERATIONS & MAINTENANCE 304.0 - - 1339 - 1339 -g

IOTV PROPELLANTS CONS'I'RlJCTION 7.2 3.3 - 32 - 0.6 32 15 OPERATIONS Llr MAINTENANCE 6 . 6 3.0 - 29 - 0.6 29 13

SUMMARY CONSTRUCTION 2687.7 2297.4 1340 11,640 1220 444 11,840 lQ121 OPERATIONS & MAINTENANCE 2490 4 -P-.--L. 20.14.tj 3694 10971 3660 396 10,971 9,008

TOTAl. 5178.1 4342. 2 5034 22p1 1 4H80 840 22,811 19,129 VEllIC1.E FIJXT

C0NSTItUCFION 14 39 12 22 110 --A-- Ol'l~llA'I'IONS k LIAINTF:NANCI: - 37 37 :17 2n IOn -_ __- _-___.

TWAI, !i I 7(i *I!) 4 2 2 I 0

4.3.2 HEAVY-LIFT LAUNCH VEHICLE (HLLV)

The primary driver in establishing HLLV requirements is the construction mass to orbit. Other factors include propellant cost/availability and envir- onmental suitability. As stated previously, an interim STS-derived HLLV will be required to satisfy SPS precursor operations (schedule limited)' and, because of its similarity to the PLV, will be defined along with that vehicle, Basic HLLV requirements are summarized in Table 4.3-4.

Table 4.3-4. HLLV Sizing-Ground Rules/Assumptions

~ . TWO-STAGE MRTICAL TAKEOFF/HORIZONTAL LANDING (VTOMLI

. FLY BACK CAPABILITY BOTH STAGES - ARES FIRST STAGE OMY

. PARALLR BURN WITH PROPEUANT CROSSFEED

l LOXIRP FIRST STAGE - LOX+ SECOND STAGE

‘,!I P, GAS GENERATOR CYCLE ENGINE - FIRST STAGE (I, (VACI - 39 SEC.1

l HI P, STAGED CONlBUSTlON ENGINE - SECOND STAGE (Is (VACI . 46 SEC.1

. STAGING MLCCITY - HEAT SINK BOOSTER CCMPATIRLE

. ClRCA lpQ0 TECHNOLOGY BASE - BACIMMC WEIGHT REDUCTION DATA

l ORBITAL PARAMETERS - UII KM 0 3L#’

l PAYl.OAD CAPABILITY - ZZI x Id KG UP’45 KG DOWN

l THRUSTMEIGHT - 130 LlFlOFFl3.D MAX

. 15% WEIGHT GROWI’I ALLOWANCUD./PL AV MARGIN

The HLLV utilizes a parallel burn mode with propellant cross-feed from the first-stage tanks to the second-stage engines. The first stage employs high chamber pressure gas generator cycle LOX/HP fueled engines with LHz cooling and the second stage employs a staged combustion engine similar to the Space Shuttle main engine (SSME) which is LOXjLHz fueled.

The HLLV configuration is shown in Figure 4.3-3 in the launch configuration. Both stages have common body diameter, wing and vertical stabilizer; however, the overall length of the second stage (orbiter) is approximately 5 m greater than the first stage (booster). The vehicle gross liftoff weight (GLOW) is 15,730,OOO lb with a payload capability of 510,000 lb to the reference earth orbit. A summary weight statement is given in Table 4.3-5. The propellant weights indicated are total loaded propellant (i.e., not usable). The second-stage weight (ULOW) includes the payload weight. During the booster ascent phase, the second-stage LOX/LH2 propellants are crossfed from the booster to achieve the parallel burn mode. Approximately 1.6 million pounds of propellant are crossfed from the booster to the orbiter during ascent.

The HLLV booster is shown in the landing configuration in Figure 4.3-4. The vehicle is approximately 300 feet in length with a wing span of 184 feet and a maximum clearance height of 116 feet. The nominal body diameter is 40 feet. The vehicle has a dry weight of 1,045,500 lb. Seven rocket engines are mounted in the aft fuselage with a nominal seal-level thrust of 2.3 million pounds each. Eight turbojet engines are mounted on the upper portion of the aft fuselage with a nominal thrust of 20,000 lb each. A detailed weight statement is given in Table 4.3-6; the vehicle propellant weight summary is projected in Table 4.3-7.

4-6

Figure 4.3-3. Reference HLLV Launch Configuration

Table 4.3-5. HLLV Mass Properties (U06)

I !s lb - GLOW 7.14 15.73

BLOW 4.92 10.84

WP2 4.49 9.89

ULOW 2.22 4.89

WP2 1.66 3.65

Payload 0.23 0.51

4-7

.

*CROSS FEED, DUAL DELTA DRY WING, L/D -7.5

RP.1 TANK ROCKET ENGINES. 7 REQ’D TOTAL THRUST. 71.441.9~33 N(S.L,

AIR BREATHER FLYBACK ENGINES - 8 fwaa

Figure 4 - 3-4 - HLLV First Stage (Booster) - Landing Configuration

Table 4.3-6. HLLV Weight Statement kgxlO-’ (IbxlO-')

SUBSYSTEn ZNO STAGE

FUSELAGE 103.41 (227.98) WING 39.20 ( 86.41) VERT:CAL TAIL 5.70 ( 12.5;) CANARD 1.39 ( 3.07) TPS 52.59 (115.94) CREU COHPARTHENT 12.70 ( 28.00) AVIONICS 3.86 ( 8.50) PERSONNEL 1.36 ( 3.001 ENV I RONMENTAL 2.59 ( 5.70) PRIM POWER 5.44 ( 12.00) HYDRAULIC SYSTEH 3.86 ( 8.50) ASCENT ENGINES 26.93 ( 59.38) KS SYSTEM 9.59 ( 21.15) LAJIDING CEAM 18.38 ( 40.51) PRDPULS I ON SY STEHS l

ATTACH AND SEPARATION APU FLYBACK ENGINES s FLYBACK PROPULSION SVSTEII m SUBSYSTEM DRY WEIGHT 286.99 (632.71) CRObI-iH MARCIH (15%) 43.0s ( 94.91) TOTAL INERT WT. 330.04 (727.62)

l \NCLllDED IN FUSELAGE UEIGHT l *lTEHS INCLUDED IN SUBSYSTEKS

IS.1 STAGE

130.73 (288.22) 78.17 y;.;;;

7.11 2.21 ( 4187)

s

3.40-1 7.50) .* .a .*

67.&148.70; l *

44.99*'( 99.18) 4.59 ( 10.12) 0.91 ( 2.00)

28.55 ( 62.95) 18.39 ( 40.54) 25.76 ( 56.80)

(909.12) (136.37)

(1045.49)

4-8

Table 4.3-7. HLLV Propellant Weight Summary (X106)

USAELL

CROSSFEED

TOTAL 8liRWED

RESIDUALS

RESERVES

ncs

ON-ORBIT

DO1 L-OFF

FLY-BACK

TOTAL LOA:ED -

-I- FIRS ~~ A 9.607 I .612 7.995 0.040 0.045 0.010

-

0.187

9.889

STAGE KG

4.358 a.732 3.626 0.018

0.020

0.005

s

0.085

SECOH LB

3.481 (1.612)

5.093 0.020 0.024 0.018

0.095 0.010

STAGE KG

I.579 (0.731) 2.310

0'. 009 0.011

0.008

0.043

0.005 w

4.486 3.648 I.655

The HLLV orbiter is depicted in Figure 4.3-5. The vehicle is approximately 317 feet in length with the same wing span, vertical height, and nominal body diameter as the booster. The orbiter employs four rocket engines with a nominal sea-level thrust of 1.19 million pounds each. The orbiter makes an upowered reentry and landing.

*CROSS FEED, DUAL-DELTA DRY WING, l/D -7.5

ROCKET ENGINES -4 REQ’D TOTAL TH(AUST - 21.129.050 N C3.L.I

12.900 DIA

Figure 4.3-S. HLLV Second Stage (Orbiter) -Landing Configuration

4-9

The cargo bay is located in the mid-fuselage and has a length of approximately 90 feet. The detailed weight statement and a propellent summary for the orbiter are included in Tables 4.3-6 and 4.3-7, respectively.

The vehicle relative staging velocity is 2127 m/set (6978 ft/sec) at an altitude of 55.15 km (181,000 ft) and a first-stage burnout range of 88.7 km (48.5 nmi). The first-stage flyback range is 387 km (211.8 nmi).

4.3.3 ELECTRIC ORBITAL TRANSFER VEHICLE (EOTV)

The EOTV depicted in Figure 4.3-6 is based upon a rigid design which can accommodate two "standard" solar blanket areas of 600.m by 750 m from the MSFC/Rockwell baseline satellite concept. The commonality of the structural configuration and construction processes with the satellite design is noted. Since the thrust levels will be very low (as compared to chemical stages), the engines and power processing units are mounted in four arrays at the lower corners of the structure/solar array. Each array contains 36 thrusters; however, only 64 thrusters are required to fire simultaneously. The additional thrusters provide redundancy when one or more arrays cannot be operated due to plume impingement on the solar array. Up to 16 thrusters, utilizing stored electrical power, are used for attitude hold only during periods of occultation. The attitude determination system is the same as the SPS, mounted at the extremities of the six vertical beams. Payload attach platforms are located so that loading/ unloading operations can be conducted from "outside" the lightweight structure.

ii,yLOAli WT. - 5.IFI@ KG EOTV DRY WT. - l.lrd KG EOTV WET WT. - l.761@ KG

Figure 4.3-6. Selected EOTV Configuration

4-10

Primary EOTV requirements are summarized in Table 4.3-8. The orbital parameters are consistent with SPS requirements and the delta-V requirement is taken from previous SEP and EOTV trajectory calculations. A 0.75% delta-V margin is included in the figure given.

Table 4.3-8. EOTV Sizing Requirements

l LEO ALTITUDE - 487 101 e 31.4. INCLllUTlON

l SOLAR INERTIAL ORlEHTATlOn

l UUNCN ANY TIME OF YEAR

l 5700 H/SEC AV REQUIREMENT

. SOUR INERTIAL ATTITUDE HOLD ONLY DURING OCCULTATION PERIODS

l 50’ PLUME CLEAMNCE l NUHDLR OF THRUSTERS - HlNlHltE

l 20% SPARE THRUSTERS - FAILURES/THRUST DIFFERENTIAL

l PERFORJWJCE LOSSES DURING THRUSTING - 5t

l ACS POWER REQUIRWENT - HAXItiUH OCCULTATION PERIOD

l ACS PROPELLANT REQUIREHENTS - 100% DUTY CYCLE

-.25t MIGHT GROWTH ALLOWAWE

The solar array has a total power output of 33.5 megawatts. Line losses of 6% and an end-of-life cell degradation of 15% yield a net power to the thruster arrays of 268.1 megawatts. The power storage system is sized on the same basis as the SPS, 200 kilowatt-hours per kilogram weight.

The GaAlAs cells are assumed to be self-annealing of electron damage occurring during transit through the Van Allen belt. A lifetime degradation in performance of 15% is consistent with basic SPS criteria.

EOTV thruster characteristics are summarized in Table 4.3-9.

Table 4.3-9. EOTV Thruster Characteristics

. MXIIWH OPERATING TEHPERATURE - 19DO. K

~ . TOTAL VOLTAGE - 8300 VOLTS

. GRID VOLTAGE - 2000 VOLTS ItAXlHUH

. REAM CURRENT - la87 AHP

. SPECIFIC IHPULSE - 8213 SEC

. THWSTER DIMETER - 76 CM

. THRUST/THRUSTER - 69.7 NEWON

. NUMER OF THRUSTERS - 144 (INCLUDES 25% SPARES)

. HM!HUH 9F 64 THWJSTEPS OPERABLE SI.W:TAI~EOUSLY

The EOTV weight and performance summary is presented in Table 4.3-10. The transfer propellant weight of 666,660 kg is the maximum that can be con- sumed by the thrusters during the transit time of 1290 days up (100 days thrusting) and the resulting return trip time of approximately 30 days (22 days thrusting).

4-11

Table 4.3-10. EOTV Weight/Performance Summary (kg)

SOLAR ARMY CELLS/STRUCTURE PObER CONOlTlONlNC

THRUSTER ARMY (4) THRUSTERS/STRUCTURE CONDUCTORS BEAlwCInML~ PROPELLANT T&WCS

ATTITUDE CONTROL SYSTEH POWER SUPPLY SYSTEM COMPONENTS PROPELLANT TAJWS

588.196

96.685

106.872

EOTV INERT UEIGHT 25% GROWTH TOTAL INERT UElCltT PROPELLANT WEIGHT

TRANSFER PROPELMT ACS PROPELLANT

EOTV LOADED WEIGHT PAYLOAD WEIGHT LEO JEPARTURE MIGHT PROPELLANT COST DELIVERED (t/KG P/L)

871,753 217,938

I .089,691 666,660

1.756.351 5.171.318

The EOTV dry weight (including growth) is approximately 1.09 10 kg and has a payload delivery capability to GE0 of 5.17x106 kg with a 10% return payload capability to LEO.

The estimated cost of $4.72/kg-payload reflects propellant costs only delivered to LEO.

4.3.4 PERSONNEL TRANSFER VEHICLE (PTV)/STS-DERIVED HLLV

The PLV and STS-derived HLLV are growth versions of the Shuttle transportation system (STS). The growth version of the PLV, Figure 4.3-7, is achieved by replacing the existing recoverable solid rocket boosters (SRB) with a pair of recoverable liquid rocket boosters (LRB). The existing orbiter and external tank are used in their current configuration. The added performance afforded by the LRB increases the orbiter payload capability to the reference STS orbit by approximately 54%, or a total capability of 45,350 kg (100,000 lb).

The STS-HLLV (Figure 4.3-8),employed in the precursor phase of SPS, is derived by replacing the STS orbiter on the PLV with a payload module and a reusable propulsion and avionics module (PAM) to provide the required orbiter functions. The PAM may be recovered ballistically or, preferably, as a down payload for the PLV. These modifications yield an STS-HLLV with a payload capability of approximately 100,000 kg.

The LRB has a gross weight of 395,000 kg, made up of 324,000 kg of propellant (278,000 kg of LO2 and 46,000 kg of LHz), and 71,000 kg of inert weight. The overall length of the LRB is 47.55 meters with a nominal diameter of 6.1 meters.

4-12

LAlJttCH CONFlGURATlON

PAVLOAO l IOOK LB tLOY‘= 3.6761 LO

8oosTm (EACH):

Glass UT n 87lU LI PROP. UT - 7lSK 18 IHERl YT = 1561: LB

I56 FT. -1 t0.U Fl WA

1

\LANOIffi ROCKETS

\ KS

PARACWE STOYAGE' ~EHGINE COVER (OPEN)

Figure 4.3-7. L02/LH2 SSME Integral Twin Ballistic Booster

REUSABLE ENGINE PO0

\

LIFTOFF WEIGHTS (lo3 kg)

PAYLOAD 100.0 EXTENAL TANK 738.3 LRB (2) 790.0 REUSABLE POD 13.7

TOTAL 1642.0

Fi'gure 4.3-8. STS HLLV Configuration

The LRB utilizes a derivative of the Space Shuttle main engine (SSME). The only difference between the LRB engines and the SSME is in nozzle expansion ratio, 35 in lieu of 77.5 to 1. The SSME-35 and its characteristics are depicted in Figure 4.3-9.

THRUST, LBF 459.000 fS.L.1 503.000 (VAC.1

EXPANSION AREA RATIO 35:l

CHAMBER PRESSURE. PSIA 3230

MIXTURE RATIO 6.0:1

SPECIFIC IMPULSE, SECONOS 406 (S.L. I US WAC.)

ENGINE WEIGHT. LBF 6340

SERVICE LIFE. HOURS 7.5 STARTS 55

ENVELOPE: LENGTH. INCHES 146 DIAMETER. INCHES

POWERHEAD 105 NOZZLE EXIT 63

Figure 4.3-g. Liquid Rocket Booster Main Engine (SSME-35)

4.3.5 PERSONNEL ORBITAL TRANSFER VEHICLE (POTV)

The POTV is the propulsive element used to transfer the personnel module (PM) from LEO to GE0 and return. The POTV concept uses a single stage to transport the PM and its crew and passengers to GEO. After initial delivery of the POTV to LEO by the STS or SPS-HLLV, the propulsive stage is subsequently refueled in LEO (at the LEO station) with sufficient propellants to execute the transfer of the PM to GEO. At GEO, the stage is refueled for a return trip of crew and passengers to LEO. The HLLV delivers crew consumables and POTV propellants to LEO and the EOTV delivers the same items required in GEO. The PM with crew/ personnel is delivered to LEO by the PLV.

The POTV configuration is shown in Figure 4.3-10, and a weight summary is given in Table 4.3-11.

The POTV utilizes two advanced space engines whose characteristics are given in Figure 4.3-11 and Table 4.3-12.

Since the POTV concept utilizes an on-orbit maintenance/refueling approach, an on-board system capable of identifying/correcting potential subsystem prob- lems in order to minimize/eliminate on-orbit checkout operations is required.

4-15

2ASE ENGINES

a-

’ 60 MAN CREW MODULE 16,000 KG

l SINGLE STAGE 01’4 36,000 KG (GE0 REFUELING)

. BOTH ELEMENTS CAPABLE OF GROWTH STS LAUNCH

Figure 4.3-10. POTV Configuration

Table 4.3-11. POTV Weight Summary

Subsystem Weight (kg) .--~

Tank (5) 1,620 Structures and lines 702 Docking ring 100 Engine (2) 490 Attitude control 235 Other 262

Subtotal 3,409 Growth (1OZ) 341

Total inert 3,750

Propellant 32,750

Total loaded 36,000 ~~__~ -

4-16

THRUST (LB)

CHAMBER PRESSURE (PSIA)

EXPANSION RATIO

MIXTURE. RATIO

SPECIFIC IMPULSE (SEC)

DIAMETER (IN.)

LENGTH (IN.)

NOZZLE RETRACTED

NOZZLE EXTENDED

Figure 4.3-11. Advanced Space Engine

Table 4.3-12. Current ASE Engine Weight

20,000

2000

400

6.0

473.0

48.5

50.5

94.0

I Fuel boost and main pumps 74.5 Oxidizer boost and main pumps 89.8 Preburner 12.4 Ducting 25.0 Combustion chamber assembly 62.8 Regen. cooled nozzle (e= 175:l) 58.4 Extendable nozzle and actuators (E = 4OO:l) 122.0 Ignition system 6.1 Controls, valves, and actuators 74.0 Heat exchanger 14.0

Total (lb)* 539.0

I3 *Based on major component current measured weights.

4-17

4.3.6 PERSONNEL MODULE (PM)

A construction sequence has been developed which requires a crew rotation every 90 days for crew complements in multiples of 60. The PM is synthesized on this basis. A limitation on PM size is established to assure compatibility with the PLV cargo bay dimensions and payload weight capacity (i.e., 4.5 m 17 m and 45,000 kg).

The PM shown in Figure 4.3-10 assumes a command station to monitor and control POTV/PM functions during flight. This function is provided in the forward section of the PM as shown. Spacing and layout of the PM is comparable to current commercial airline practice. Seating is provided on the basis of one meter, front to rear, and a width of 0.72 meter. .PM mass was established on the basis of 110 kg/man (including personal effects) and approximately 190 kg/man for module mass. The PM design has provisions for 60 passengers and two flight crew members.

4.3.7 INTRA-ORBIT TRANSFER VEHICLE (IOTV)

On-orbit mobility systems are syntehsized in terms of application and concept only. On-orbit elements considered here are powered by a chemical (LOX/LH ) propulsion system. At least three distinct applications have been identified: (1) the need to transfer cargo from the HLLV to the EOTV in LEO, and from the EOTV to the SPS construction base in GEO; (2) the need to move materials about the SPS construction base; and (3) the probable need to move men or materials between operational SPS's. A "free-flyer" teleoperator concept is assumed.

Sizing of the IOTV is based on a minimum safe separation distance between EOTV and the SPS base of 10 km. The assumed transfer time is in the order of two hours (round trip), which equates to a AV requirement on the order of 3 to 5 misec. A single advanced space engine (ASE) is employed with a specific impulse of 473 set (see Section 4.3.5 for complete engine description). The pertinent IOTV parameters are summarized in Table 4.3-13.

Table 4.3-13. IOTV Weight Summary

SUBSYSTEM WE I GHT (kg)

ENGINE (1 ASE) 245 PROPELLANT TANKS IS STRUCTURE AND LINES s 15 DOCKING RI tJG 100 All- I TUDE CONTROL 50 OTHER 100 SUBTOTAL 525 GROWTH (10%) 53 TOTAL I NE RT 57 PROPELLANT 30:: TOTAL LOADED 878

4-18

4.4 LEO OPERATIONAL BASE

(TBD)

4.5 CARGO AND PERSONNEL LAUNCH AND RECOVERY FACILITIES

(TBD)

4.6 BASE SUPPORT FACILITIES

(TBD)

4.7 LOGISTIC FACILITIES

(TBD)

4.8 SPS GROUND RECTENNA FACILITIES

(TBD)

4-19

1. Report No. NASA CR-3324 4. Title and Subwle

2. Government Accuion No.

SATELLITE POWER SYSTEMS (SPS) CONCEPT DEFINITION STUDY VOLUME VII - SYSTm/SUBSYSTEM REQUIREMENTS DATA BOOK

7. Author(s)

G. M. Hanley

9. Performing Organization Name and Address

Rockwell International 12214 La kewood Boul evard Downey, CA 90241

2. Sponsoring Agency Name and Address

Notional Aeronautics and Space Administration Washington, D. C. 20546 5. Surqalemcntary Notes

Marshall Technical Monitor: C. H. Guttman Volume VII of Final Report

3. Recipient’8 Catalog No.

5. Repon Date September 1980

6. Performing Organization Code

6. PerforminO Organization Repon No.

SSD 79-0010-7 10. Work Unit No.

11. Concracc or Grant No

NAS8-32475 13. Type of Rcpon and Period Gwercd

Contractor Report 14. Sponsoring Agency code

This volume of the Satellite Power Systems (SPS) Concept Definition Study final report summarizes the basic requirements used as a guide to systems analysis and is a basis for the selection of candidate SPS point design(s). Initially, these collected data reflected the level of definition resulting from the evaluation of a broad spectrum of SPS concepts. As the various concepts matured these requirements were update to reflect the requirements identified for the projected satellite system/subsys tern point design(s). Earlier studies (reported in Volumes I -VII, SD 79-AP-0023, dated April 1978) established two candidate conce] which were presented to the NASA for consideration. NASA, in turn, util: these and other concepts developed under the auspices of other contracts, established a baseline or reference concept which was to be the basis fol future evaluation and point design. This volume defines the identified subsystem/systems requirenents, and where appropriate, presents recommendations for alternate approaches which may represent improved design features.

‘. Kav Words &q9ested by Author(r)) 16. DMribution Statement

Sys tern Requirements Subsys tern Requirements

Unclassified - Unlimited

Sys tern Description Subsys tern Description Satellite Power System I Subject Category 44

1 kcurucy Oawf. (of this report1 20. Securlcy Clawf. (of this pagel 21. No. of Pw 22. Price Unclassified Unclassified 128 A07

For sale by the Nafional TechnIcal Infocmatlon Service. Sprqfleld. Vlrglnla 22161 NASA-Lang1 ey > 1980

NASA - National Space Society · ILLUSTRATIONS Figure 1.1-l 1.1-2 1.2-l 1.2-2 1.2-3 1.2-4 1.2-5 1.2-6 1.2-7 1.2-8 1.2-9 1.2-10 1.2-11 1.2-12 1.2-13 1.2-14 1.2-15 1.3-l 2.1-l 2.1-2

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